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Sample records for rocket solid propellants

  1. Some typical solid propellant rocket motors

    NARCIS (Netherlands)

    Zandbergen, B.T.C.

    2013-01-01

    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  2. Rocket Solid Propellant Alternative Based on Ammonium Dinitramide

    Directory of Open Access Journals (Sweden)

    Grigore CICAN

    2017-03-01

    Full Text Available Due to the continuous run for a green environment the current article proposes a new type of solid propellant based on the fairly new synthesized oxidizer, ammonium dinitramide (ADN. Apart of having a higher specific impulse than the worldwide renowned oxidizer, ammonium perchlorate, ADN has the advantage, of leaving behind only nitrogen, oxygen and water after decomposing at high temperatures and therefore totally avoiding the formation of hydrogen chloride fumes. Based on the oxidizer to fuel ratios of the current formulations of the major rocket solid booster (e.g. Space Shuttle’s SRB, Ariane 5’s SRB which comprises mass variations of ammonium perchlorate oxidizer (70-75%, atomized aluminum powder (10-18% and polybutadiene binder (12-20% a new solid propellant was formulated. As previously stated, the new propellant formula and its variations use ADN as oxidizer and erythritol tetranitrate as fuel, keeping the same polybutadiene as binder.

  3. SRM (Solid Rocket Motor) propellant and polymer materials structural modeling

    Science.gov (United States)

    Moore, Carleton J.

    1988-01-01

    The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

  4. Estimation of Pressure Index and Temperature Sensitivity Coefficient of Solid Rocket Propellants by Static Evaluation

    OpenAIRE

    Himanshu Shekhar

    2009-01-01

    Burning rate of a solid rocket propellant depends on pressure and temperature. Conventional strand burner and Crawford bomb test on propellant strands was conducted to assess these dependent parameters. However, behaviour of propellant in rocket motor is different from its behaviour in strand form. To overcome this anomaly, data from static evaluation of rocket motor was directly used for assessment of these burningrate controlling parameters. The conventional empirical power law (r=aoexp[p{T...

  5. EXTRUSION FORMING OF A DOUBLE BASE SOLID ROCKET PROPELLANT BY FINITE ELEMENT METHOD

    Directory of Open Access Journals (Sweden)

    Barış KALAYCIOĞLU

    2007-01-01

    Full Text Available In this study, three dimensional modelling of extrusion forming of a double base solid rocket propellant is performed on Ansys® finite element simulation package. For the purpose of initial model construction and later comparisons with elastoviscoplastik model, the solid propellant is assumed to obey the elastic-plastic material response during the direct extrusion process. Taking into account the contact surface behavior with Coulomb friction and geometric and material nonlinearities, an incremental large large strain solution methodology has been adapted in the simulation. The hydrostatic pressure, stress, strain, and displacement values during extrusion of the solid rocket propellant are obtained from the simulation.

  6. Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

    Science.gov (United States)

    Gill, W.; Cruz-Cabrera, A. A.; Donaldson, A. B.; Lim, J.; Sivathanu, Y.; Bystrom, E.; Haug, A.; Sharp, L.; Surmick, D. M.

    2014-11-01

    Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

  7. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Directory of Open Access Journals (Sweden)

    Zhukov Ilya S.

    2016-01-01

    Full Text Available On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  8. Development of high temperature materials for solid propellant rocket nozzle applications

    Science.gov (United States)

    Manning, C. R., Jr.; Lineback, L. D.

    1974-01-01

    Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

  9. Regarding the evaluation of the solid rocket propellant response function to pressure coupling

    Directory of Open Access Journals (Sweden)

    Ioan ION

    2011-03-01

    Full Text Available High frequency combustion instabilities imply a major risk for the solid rocket motor stableworking and they are directly linked to the propellant response to chamber pressure coupling. Thisarticle discusses a laboratory testing method for the measurement and evaluation of the pressurecoupled response for non-metalized propellants in a first stage. Experimental researches were donewith an adequate setup, built and improved in our lab, able to evaluate the propellant response byinterpreting the pressure oscillations damping in terms of propellant response. Our paper aims atdefining a linearized one-dimensional flow study model to analyze the disturbed operation of the solidpropellant rocket motors. Based on the applied model we can assert that the real part of propellantresponse is a function of the oscillations damping, acoustic energy in the motor chamber and variouslosses in the burning chamber. The imaginary part of propellant response mainly depends on thenormalized pulsation, on the burning chamber gas column and on the pressure oscillations frequency.Our research purpose was obviously to minimize the risk of the combustion instabilities effects on therocket motors working, by experimental investigations using jet modulating techniques and sustainedby an interesting study model based on the perturbation method.

  10. Neki aspekti performansi raketnih goriva i pravci razvoja čvrstih raketnih goriva / Some aspects of rocket propellant performances and solid rocket propellant development tendencies

    Directory of Open Access Journals (Sweden)

    Radun Jeremić

    2004-01-01

    Full Text Available Radi povećanja korisnog tereta raketnih sistema napori istraživača u oblasti raketnih goriva usmereni su, pre svega, na osvajanje novih sastava sa visokim sadržajem energije, koji će obezbediti veći zapreminski specifični impuls. U radu je dat kratak pregled dosadašnjeg razvoja raketnih goriva, sa posebnim osvrtom na najnovija dostignuća u poboljšanju energetskih karakteristika postojećih sastava, kao i osvajanje potpuno novih sastava visokoenergetskih čvrstih raketnih goriva. / In order to increase pay loads in missile systems, research work in the domain of rocket propellants concentrates on conquering new, high-energy compositions providing a greater volume specific impulse. A short survey of rocket propellant development up to present days is followed by the most recent achievements in enhancing energy characteristics of existing compositions and in conquering completely new compositions of high-energy solid rocket propellants.

  11. Finite element analysis of propellant of solid rocket motor during ship motion

    Directory of Open Access Journals (Sweden)

    Kai Qu

    2013-03-01

    Full Text Available In order to simulate the stress and strain of solid rocket motors (SRMs, a finite element analysis model was established. The stress spectra of the SRM elements with respect to time in the case that the vessel cruises under a certain shipping condition were obtained by simulation. According to the analysis of the simulation results, a critical zone was confirmed, and the Mises stress amplitudes of the different critical zones were acquired. The results show that the maximum stress and strain of SRM are less than the maximum tensile strength and elongation, respectively, of the propellant. The cumulative damage of the motor must also be evaluated by random fatigue loading.

  12. Investigation of the flow turning loss in unstable solid propellant rocket motors

    Science.gov (United States)

    Matta, Lawrence Mark

    The goal of this study was to improve the understanding of the flow turning loss, which contributes to the damping of axial acoustic instabilities in solid propellant rocket motors. This understanding is needed to develop practical methods for designing motors that do not exhibit such instabilities. The flow turning loss results from the interaction of the flow of combustion products leaving the surface of the propellant with the acoustic field in an unstable motor. While state of the art solid rocket stability models generally account for the flow turning loss, its magnitude and characteristics have never been fully investigated. This thesis describes a combined theoretical, numerical, and experimental investigation of the flow turning loss and its dependence upon various motor design and operating parameters. First, a one dimensional acoustic stability equation that verifies the existence of the flow turning loss was derived for a chamber with constant mean pressure and temperature. The theoretical development was then extended to include the effects of mean temperature gradients to accommodate combustion systems in which mean temperature gradients and heat losses are significant. These analyses provided the background and expressions necessary to guide an experimental study. The relevant equations were then solved for the developed experimental setup to predict the behavior of the flow turning loss and the other terms of the developed acoustic stability equation. This was followed by and experimental study in which the flow turning region of an unstable solid propellant rocket motor was simulated. The setup was used, with and without combustion, to determine the dependence of the flow turning loss upon operating conditions. These studies showed that the flow turning loss strongly depends upon the gas velocity at the propellant surface and the location of the flow turning region relative to the standing acoustic wave. The flow turning loss measured in the

  13. Effect of ammonium perchlorate grain sizes on the combustion of solid rocket propellant

    Energy Technology Data Exchange (ETDEWEB)

    Hegab, A.; Balabel, A. [Menoufia Univ., Menoufia (Egypt). Faculty of Engineering

    2007-07-01

    The combustion of heterogeneous solid rocket propellant consisting of ammonium perchlorate (AP) particles was discussed with reference to the chemical and physical complexity of the propellant and the microscopic scale of the combustion zone. This study considered the primary flame between the decomposition products of the binder and the AP oxidizer; the primary diffusion flame from the oxidizer; density and conductivity of the AP and binder; temperature-dependent gas-phase transport properties; and, an unsteady non-planer regression surface. Three different random packing disc models for the AP particles imbedded in a matrix of a hydroxyl terminated polybutadience (HTPB) fuel-binder were used as a base of the combustion code. The models have different AP grain sizes and distribution with the fuel binder. A 2D calculation was developed for the combustion of heterogeneous solid propellant, accounting for the gas phase physics, the solid phase physics and an unsteady non-planar description of the regressing propellant surface. The mathematical model described the unsteady burning of a heterogeneous propellant by simultaneously solving the combustion fields in the gas phase and the thermal field in the solid phase with appropriate jump condition across the gas/solid interface. The gas-phase kinetics was represented by a two-step reaction mechanism for the primary premixed flame and the primary diffusion flame between the decomposition products of the HTPB and the oxidizer. The essentially-non-oscillatory (ENO) scheme was used to describe the propagation of the unsteady non-planer regression surface. The results showed that AP particle size has a significant effect on the combustion surface deformation as well as on the burning rate. This study also determined the effect of various parameters on the surface propagation speed, flame structure, and the burning surface geometry. The speed by which the combustion surface recedes was found to depend on the exposed pressure

  14. Low toxicity rocket propellants

    NARCIS (Netherlands)

    Wink, J.

    2014-01-01

    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  15. Effect of ambient vibration on solid rocket motor grain and propellant/liner bonding interface

    Science.gov (United States)

    Cao, Yijun; Huang, Weidong; Li, Jinfei

    2017-05-01

    In order to study the condition of structural integrity in the process of the solid propellant motor launching and transporting, the stress and strain field analysis were studied on a certain type of solid propellant motor. the vibration acceleration on the solid propellant motors' transport process were monitored, then the original vibration data was eliminated the noise and the trend term efficiently, finally the characteristic frequency of vibration was got to the finite element analysis. Experiment and simulation results show that the monitored solid propellant motor mainly bear 0.2 HZ and 15 HZ low frequency vibration in the process of transportation; Under the low frequency vibration loading, solid propellant motor grain stress concentration position is respectively below the head and tail of the propellant/liner bonding surface and the grain roots.

  16. Removing hydrochloric acid exhaust products from high performance solid rocket propellant using aluminum-lithium alloy

    Energy Technology Data Exchange (ETDEWEB)

    Terry, Brandon C., E-mail: terry13@purdue.edu [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Sippel, Travis R. [Department of Mechanical Engineering, Iowa State University, 2025 Black Engineering, Ames, IA 50011 (United States); Pfeil, Mark A. [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Gunduz, I.Emre; Son, Steven F. [School of Mechanical Engineering, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States)

    2016-11-05

    Highlights: • Al-Li alloy propellant has increased ideal specific impulse over neat aluminum. • Al-Li alloy propellant has a near complete reduction in HCl acid formation. • Reduction in HCl was verified with wet bomb experiments and DSC/TGA-MS/FTIR. - Abstract: Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solid propellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (I{sub SP}). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal I{sub SP} by ∼7 s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solid propellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5 ± 4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption.

  17. Alternate Propellant Thermal Rocket Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Alternate Propellant Thermal Rocket (APTR) is a novel concept for propulsion of space exploration or orbit transfer vehicles. APTR propulsion is provided by...

  18. Removing hydrochloric acid exhaust products from high performance solid rocket propellant using aluminum-lithium alloy.

    Science.gov (United States)

    Terry, Brandon C; Sippel, Travis R; Pfeil, Mark A; Gunduz, I Emre; Son, Steven F

    2016-11-05

    Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solid propellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (ISP). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal ISP by ∼7s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solid propellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5±4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption. Copyright © 2016 Elsevier B.V. All rights reserved.

  19. Liquid propellant rocket combustion instability

    Science.gov (United States)

    Harrje, D. T.

    1972-01-01

    The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

  20. Numerical investigation of the influence of crystallization of ultrafine particles of aluminum oxide on energy characteristics of solid-propellant rocket engine

    Science.gov (United States)

    Dyachenko, N. N.; Dyachenko, L. I.

    2014-08-01

    The results of numerical investigation of a multiphase flow considering coagulation, crushing and crystallization of the particles of polydispersed condensate in the nozzles of solid-propellant rocket engine are presented. The influence of particles crystallization on the energy characteristics of the engine is shown.

  1. ISRO's solid rocket motors

    Science.gov (United States)

    Nagappa, R.; Kurup, M. R.; Muthunayagam, A. E.

    1989-08-01

    Solid rocket motors have been the mainstay of ISRO's sounding rockets and the first generation satellite launch vehicles. For the new launch vehicle under development also, the solid rocket motors contribute significantly to the vehicle's total propulsive power. The rocket motors in use and under development have been developed for a variety of applications and range in size from 30 mm dia employing 450 g of solid propellant—employed for providing a spin to the apogee motors—to the giant 2.8 m dia motor employing nearly 130 tonnes of solid propellant. The initial development, undertaken in 1967 was of small calibre motor of 75 mm dia using a double base charge. The development was essentially to understand the technological elements. Extruded aluminium tubes were used as a rocket motor casing. The fore and aft closures were machined from aluminium rods. The grain was a seven-pointed star with an enlargement of the port at the aft end and was charged into the chamber using a polyester resin system. The nozzle was a metallic heat sink type with graphite throat insert. The motor was ignited with a black powder charge and fired for 2.0 s. Subsequent to this, further developmental activities were undertaken using PVC plastisol based propellants. A class of sounding rockets ranging from 125 to 560 mm calibre were realized. These rocket motors employed improved designs and had delivered lsp ranging from 2060 to 2256 Ns/kg. Case bonding could not be adopted due to the higher cure temperatures of the plastisol propellants but improvements were made in the grain charging techniques and in the design of the igniters and the nozzle. Ablative nozzles based on asbestos phenolic and silica phenolic with graphite inserts were used. For the larger calibre rocket motors, the lsp could be improved by metallic additives. In the early 1970s designs were evolved for larger and more efficient motors. A series of 4 motors for the country's first satellite launch vehicle SLV-3 were

  2. Investigation of the effects of solid rocket motor propellant composition on plume signature

    Science.gov (United States)

    Snaza, Clay J.

    1994-06-01

    Three propellants with aluminum/silicon weight percentages of 18/0%, 13.5/4.5%, and 12/6% were fired in a subscale motor to determine if the plume infrared signature could be reduced without a significant loss in specific impulse. Spectral measurements from 2.5 to 5.5 micrometers and thermal measurements from 3.5 to 5.0 micrometers were made. Plume particle size measurements showed that only particles with small diameters (less than 1.93 micrometers) were present with any significant volume. Replacing a portion of the aluminum in a highly metallized solid propellant with silicon was found to eliminate the Al2O3 in favor of SiO2 and Al6SiO13, without any change in particulate mass concentration or any large change in particle size distribution. These particulates were found to have significantly lower absorptivity than Al2O3. An additional investigation was conducted to determine the particle size distribution at the nozzle entrance. Malvern ensemble scattering, phase-Doppler single particle scattering, and laser transmittance measurements made through windows in the combustion chamber at the nozzle entrance indicated that large particles were present (to 250 micrometers). However, most of the mass of the particles was contained in particles with diameters smaller than 5 micrometers. Approximate calculations made with the measured data showed that if 100 micrometer particles are present with the smoke (particles with diameters less than 2 micrometers) they could account for only approximately 10% of the article volume.

  3. Simple-1: Development stage of the data transmission system for a solid propellant mid-power rocket model

    Science.gov (United States)

    Yarce, Andrés; Sebastián Rodríguez, Juan; Galvez, Julián; Gómez, Alejandro; García, Manuel J.

    2017-06-01

    This paper presents the development stage of a communication module for a solid propellant mid-power rocket model. The communication module was named. Simple-1 and this work considers its design, construction and testing. A rocket model Estes Ventris Series Pro II® was modified to introduce, on the top of the payload, several sensors in a CanSat form factor. The Printed Circuit Board (PCB) was designed and fabricated from Commercial Off The Shelf (COTS) components and assembled in a cylindrical rack structure similar to this small format satellite concept. The sensors data was processed using one Arduino Mini and transmitted using a radio module to a Software Defined Radio (SDR) HackRF based platform on the ground station. The Simple-1 was tested using a drone in successive releases, reaching altitudes from 200 to 300 meters. Different kind of data, in terms of altitude, position, atmospheric pressure and vehicle temperature were successfully measured, making possible the progress to a next stage of launching and analysis.

  4. Burning mechanism of aluminized solid rocket propellants based on energetic binders

    Energy Technology Data Exchange (ETDEWEB)

    Babuk, Valery A.; Dolotkazin, Ildar N.; Glebov, Andrey A. [Baltic State Technical University (BSTU), First Krasnoarmeyskaya Str. 1, Saint Petersburg (Russian Federation)

    2005-09-01

    This paper reports results obtained from an experimental study of the combustion mechanism of aluminized propellants based on an energetic binder. The techniques used in this investigation include:. (Abstract Copyright [2005], Wiley Periodicals, Inc.)

  5. Asbestos Free Insulation Development for the Space Shuttle Solid Propellant Rocket Motor (RSRM)

    Science.gov (United States)

    Allred, Larry D.; Eddy, Norman F.; McCool, A. A. (Technical Monitor)

    2000-01-01

    Asbestos has been used for many years as an ablation inhibitor in insulating materials. It has been a constituent of the AS/NBR insulation used to protect the steel case of the RSRM (Reusable Solid Rocket Motor) since its inception. This paper discusses the development of a potential replacement RSRM insulation design, several of the numerous design issues that were worked and processing problems that were resolved. The earlier design demonstration on FSM-5 (Flight Support Motor) of the selected 7% and 11% Kevlar(registered) filled EPDM (KF/EPDM) candidate materials was expanded. Full-scale process simulation articles were built and FSM-8 was manufactured using multiple Asbestos Free (AF) components and materials. Two major problems had to be overcome in developing the AF design. First, bondline corrosion, which occurred in the double-cured region of the aft dome, had to be eliminated. Second, KF/EPDM creates high levels of electrostatic energy (ESE), which does not readily dissipate from the insulation surface. An uncontrolled electrostatic discharge (ESD) of this surface energy during many phases of production could create serious safety hazards. Numerous processing changes were implemented and a conductive paint was developed to prevent exposed external insulation surfaces from generating ESE/ESD. Additionally, special internal instrumentation was incorporated into FSM-8 to record real-time internal motor environment data. These data included inhibitor insulation erosion rates and internal thermal environments. The FSM-8 static test was successfully conducted in February 2000 and much valuable data were obtained to characterize the AF insulation design.

  6. Nanoparticles for solid rocket propulsion

    Energy Technology Data Exchange (ETDEWEB)

    Galfetti, L [Politecnico di Milano, SPLab, Milan (Italy); De Luca, L T [Politecnico di Milano, SPLab, Milan (Italy); Severini, F [Politecnico di Milano, SPLab, Milan (Italy); Meda, L [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marra, G [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marchetti, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Regi, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Bellucci, S [INFN, Laboratori Nazionali di Frascati, Frascati (Italy)

    2006-08-23

    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

  7. Integrated model of a composite propellant rocket

    Science.gov (United States)

    Miccio, Francesco

    2016-12-01

    The combustion of composite solid propellants was investigated and an available numerical model was improved for taking into account the change of pressure, when the process occurs in a confined environment, as inside a rocket. The pressure increase upon ignition is correctly described by the improved model for both sandwich and dispersed particles propellants. In the latter case, self-induced fluctuations in the pressure and in all other computed variables occur, as consequence of the periodic rise and depletion of oxidizer particles from the binder matrix. The comparison with the results of the constant pressure model shows a different fluctuating profile of gas velocity, with a possible second order effect induced by the pressure fluctuations.

  8. Enantioselective synthesis of aziridines using asymmetric transfer hydrogenation as a precursor for chiral derivatives used as bonding agent for rocket solid propellants

    Directory of Open Access Journals (Sweden)

    Aparecida M. Kawamoto

    2002-11-01

    Full Text Available A rapid, expedient and enantioselective method for the synthesis of beta-hydroxy amines and monosubstituted aziridines in up to 99% e.e., via asymmetric transfer hydrogenation of a-amino ketones and cyclisation through treatment with tosyl chloride and base, is described. (1R,2R-N-(para-toluenesulfonyl-1,2-ethylenediamine with formic acid has been utilised as a ligand for the Ruthenium (II catalysed enantioselective transfer hydrogenation of the ketones.The chiral 2-methyl aziridine, which is a potentially more efficient bonding agent for Rocket Solid Propellant has been successfully achieved.

  9. VIABILITY OF BACILLUS SUBTILIS SPORES IN ROCKET PROPELLANTS.

    Science.gov (United States)

    GODDING, R M; LYNCH, V H

    1965-01-01

    The sporicidal activity of components used in liquid and solid rocket propellants was tested by use of spores of Bacillus subtilis dried on powdered glass. Liquid propellant ingredients tested were N(2)O(4), monomethylhydrazine and 1,1-dimethylhydrazine. N(2)O(4) was immediately sporicidal; the hydrazines were effective within several days. Solid propellants consisted of ammonium perchlorate in combination with epoxy resin (EPON 828), tris-1-(2-methyl) aziridinyl phosphine oxide, bis-1-(2-methyl) aziridinyl phenylphosphine oxide, and three modified polybutadiene polymers. There was no indication of appreciable sporicidal activity of these components.

  10. Dynamic mechanical analysis of double base rocket propellants

    Directory of Open Access Journals (Sweden)

    Marcin Cegła

    2016-03-01

    Full Text Available The article presents dynamic mechanical analysis (DMA for solid rocket propellants testing. Principles of operation and measured values are briefly described. The authors refer to the previous research of PTFE material and literature data providing information about proper experimental conditions and influence of measurement frequency, load amplitude, and heating rate on the results of DMA tests. The experimental results of solid double-base rocket propellant testing obtained on the N Netzsch DMA 242 device are presented. Mechanical properties such as the dynamic storage modulus E´, the dynamic loss modulus E˝ and tan(δ were measured within temperature range from (–120°C to (+90°C at the heating rate of 1 K/min. The test sample was subjected to a dual cantilever multi-frequency test. Special attention was paid to determination of the glass transition temperature of the tested propellant in reference to the NATO standardization agreement 4540 as well as influence of the measurement frequency on the glass transition.[b]Keywords[/b]: Dynamic mechanical analysis, solid rocket propellants, glass transition temperature

  11. The Advanced Solid Rocket Motor

    Science.gov (United States)

    Mitchell, Royce E.

    1992-08-01

    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  12. Nano-aluminum as energetic material for rocket propellants

    Energy Technology Data Exchange (ETDEWEB)

    Meda, L. [Istituto G. Donegani, Polimeri Europa, 28100 Novara (Italy)], E-mail: laura.meda@polimerieuropa.com; Marra, G. [Istituto G. Donegani, Polimeri Europa, 28100 Novara (Italy); Galfetti, L.; Severini, F.; De Luca, L. [Politecnico di Milano, Solid Propulsion Laboratory (SP Lab), 20158 Milano (Italy)

    2007-09-15

    A characterization of differently sized aluminum powders, by using BET (specific surface measurements), Scanning Electron Microscopy (SEM), X-Ray Diffraction (XRD), and X-ray Photoelectron Spectroscopy (XPS), was performed in order to evaluate their performance in solid propellant. These aluminum powders were used in manufacturing composite rocket propellants, that are based on Ammonium Perchlorate (AP) as oxidizer and Hydroxyl-Terminated-PolyButadiene (HTPB) as binder. The reference formulation was AP/HTPB/Al with 68/17/15 mass fractions, respectively. The ballistic characterization of studied propellants, made in terms of steady burning rates, showed how better is the performance of nano-aluminized compared to micro-aluminized propellants. Measurements of Al powder ignition time and temperature were also carried out.

  13. Thiokol Solid Rocket Motors

    Science.gov (United States)

    Graves, S. R.

    2000-01-01

    This paper presents viewgraphs on thiokol solid rocket motors. The topics include: 1) Communications; 2) Military and government intelligence; 3) Positioning satellites; 4) Remote sensing; 5) Space burial; 6) Science; 7) Space manufacturing; 8) Advertising; 9) Space rescue space debris management; 10) Space tourism; 11) Space settlements; 12) Hazardous waste disposal; 13) Extraterrestrial resources; 14) Fast package delivery; and 15) Space utilities.

  14. A research on polyether glycol replaced APCP rocket propellant

    Science.gov (United States)

    Lou, Tianyou; Bao, Chun Jia; Wang, Yiyang

    2017-08-01

    Ammonium perchlorate composite propellant (APCP) is a modern solid rocket propellant used in rocket vehicles. It differs from many traditional solid rocket propellants by the nature of how it is processed. APCP is cast into shape, as opposed to powder pressing it with black powder. This provides manufacturing regularity and repeatability, which are necessary requirements for use in the aerospace industry. For traditional APCP, ingredients normally used are ammonium peroxide, aluminum, Hydroxyl-terminated polybutadiene(HTPB), curing agency and other additives, the greatest disadvantage is that the fuel is too expensive. According to the price we collected in our country, a single kilogram of this fuel will cost 200 Yuan, which is about 35 dollars, for a fan who may use tons of the fuel in a single year, it definitely is a great deal of money. For this reason, we invented a new kind of APCP fuel. Changing adhesive agency from cross-linked htpb to cross linked polyether glycol gives a similar specific thrust, density and mechanical property while costs a lower price.

  15. Fuel-Cell Power Source Based on Onboard Rocket Propellants

    Science.gov (United States)

    Ganapathi, Gani; Narayan, Sri

    2010-01-01

    The use of onboard rocket propellants (dense liquids at room temperature) in place of conventional cryogenic fuel-cell reactants (hydrogen and oxygen) eliminates the mass penalties associated with cryocooling and boil-off. The high energy content and density of the rocket propellants will also require no additional chemical processing. For a 30-day mission on the Moon that requires a continuous 100 watts of power, the reactant mass and volume would be reduced by 15 and 50 percent, respectively, even without accounting for boiloff losses. The savings increase further with increasing transit times. A high-temperature, solid oxide, electrolyte-based fuel-cell configuration, that can rapidly combine rocket propellants - both monopropellant system with hydrazine and bi-propellant systems such as monomethyl hydrazine/ unsymmetrical dimethyl hydrazine (MMH/UDMH) and nitrogen tetroxide (NTO) to produce electrical energy - overcomes the severe drawbacks of earlier attempts in 1963-1967 of using fuel reforming and aqueous media. The electrical energy available from such a fuel cell operating at 60-percent efficiency is estimated to be 1,500 Wh/kg of reactants. The proposed use of zirconia-based oxide electrolyte at 800-1,000 C will permit continuous operation, very high power densities, and substantially increased efficiency of conversion over any of the earlier attempts. The solid oxide fuel cell is also tolerant to a wide range of environmental temperatures. Such a system is built for easy refueling for exploration missions and for the ability to turn on after several years of transit. Specific examples of future missions are in-situ landers on Europa and Titan that will face extreme radiation and temperature environments, flyby missions to Saturn, and landed missions on the Moon with 14 day/night cycles.

  16. Closed-cycle liquid propellant rocket engines

    Science.gov (United States)

    Kuznetsov, N. D.

    1993-06-01

    The paper presents experience gained by SSSPE TRUD in development of NK-33, NK-43, NK-39, and NK-31 liquid propellant rocket engines, which are reusable, closed-cycle type, working on liquid oxygen and kerosene. Results are presented showing the engine structure efficiency, configuration rationality, and optimal thrust values which provide the following specific parameters: specific vacuum impulses in the range 331-353 s (for NK-33 and NK-31 engines, respectively) and specific weight of about 8 kg/tf (NK-33 and NK-43 engines). The problems which occurred during engine development and the study of the main components of these engines are discussed. The important technical data, materials, methodology, and bench development data are presented for the gas generator, turbopump assembly, combustion chamber and full-scale engines.

  17. Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

    Data.gov (United States)

    National Aeronautics and Space Administration — A dynamical model describing the multi-stage process of rocket propellant loading has been developed. It accounts for both the nominal and faulty regimes of...

  18. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Spark ignition of a bi-propellant rocket engine is a classic, proven, and generally reliable process. However, timing can be critical, and the control logic,...

  19. On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

    Science.gov (United States)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

  20. Solid rocket technology advancements for space tug and IUS applications

    Science.gov (United States)

    Ascher, W.; Bailey, R. L.; Behm, J. W.; Gin, W.

    1975-01-01

    In order for the shuttle tug or interim upper stage (IUS) to capture all the missions in the current mission model for the tug and the IUS, an auxiliary or kick stage, using a solid propellant rocket motor, is required. Two solid propellant rocket motor technology concepts are described. One concept, called the 'advanced propulsion module' motor, is an 1800-kg, high-mass-fraction motor, which is single-burn and contains Class 2 propellent. The other concept, called the high energy upper stage restartable solid, is a two-burn (stop-restartable on command) motor which at present contains 1400 kg of Class 7 propellant. The details and status of the motor design and component and motor test results to date are presented, along with the schedule for future work.

  1. Solid Rocket Launch Vehicle Explosion Environments

    Science.gov (United States)

    Richardson, E. H.; Blackwood, J. M.; Hays, M. J.; Skinner, T.

    2014-01-01

    Empirical explosion data from full scale solid rocket launch vehicle accidents and tests were collected from all available literature from the 1950s to the present. In general data included peak blast overpressure, blast impulse, fragment size, fragment speed, and fragment dispersion. Most propellants were 1.1 explosives but a few were 1.3. Oftentimes the data from a single accident was disjointed and/or missing key aspects. Despite this fact, once the data as a whole was digitized, categorized, and plotted clear trends appeared. Particular emphasis was placed on tests or accidents that would be applicable to scenarios from which a crew might need to escape. Therefore, such tests where a large quantity of high explosive was used to initiate the solid rocket explosion were differentiated. Also, high speed ground impacts or tests used to simulate such were also culled. It was found that the explosions from all accidents and applicable tests could be described using only the pressurized gas energy stored in the chamber at the time of failure. Additionally, fragmentation trends were produced. Only one accident mentioned the elusive "small" propellant fragments, but upon further analysis it was found that these were most likely produced as secondary fragments when larger primary fragments impacted the ground. Finally, a brief discussion of how this data is used in a new launch vehicle explosion model for improving crew/payload survival is presented.

  2. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    Directory of Open Access Journals (Sweden)

    Abdelaziz Almostafa

    2018-01-01

    Full Text Available Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning, erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameters affect on erosive burning. Investigate the phenomena of the erosive burning by using the 2’inch rocket motor and modified one. Different tests applied to fulfil all the parameters that calculated out from the experiments and by studying the pressure time curve and erosive burning phenomena.

  3. Project SQUID. Liquid Propellant Rockets. Volume 2, Part 2

    Science.gov (United States)

    1947-06-30

    problem-, directed tow ard droppable .JATO unit is being developed for the XB-45 the desiga or levelopment of a specific power plant. airplane. This...operations such as propellant dynamometer stand and absorb the thrust of the analysis and instrument calibration. rocket. In addition to these main

  4. Ultrasonic investigation of mechanical properties of double base rocket propellants

    NARCIS (Netherlands)

    Schroeff, J.A. van der; Boer, R.S. de

    1976-01-01

    For a series of double base rocket propellants and for poly-methylmethacrylate (PMMA) the longitudinal and transverse sound wave velocities are measured at a frequency of 0.351 MHz in t h e temperature range of −40°C to +60°C. The relations between these acoustic properties and mechanical properties

  5. Metallic Hydrogen: A Game Changing Rocket Propellant

    Science.gov (United States)

    Silvera, Isaac F.

    2016-01-01

    The objective of this research is to produce metallic hydrogen in the laboratory using an innovative approach, and to study its metastability properties. Current theoretical and experimental considerations expect that extremely high pressures of order 4-6 megabar are required to transform molecular hydrogen to the metallic phase. When metallic hydrogen is produced in the laboratory it will be extremely important to determine if it is metastable at modest temperatures, i.e. remains metallic when the pressure is released. Then it could be used as the most powerful chemical rocket fuel that exists and revolutionize rocketry, allowing single-stage rockets to enter orbit and chemically fueled rockets to explore our solar system.

  6. Solid rocket technology advancement for Space Tug and IUS applications. [Interim Upper Stage

    Science.gov (United States)

    Ascher, W.; Bailey, R. L.; Behm, J. W.; Gin, W.

    1975-01-01

    Two-burn restartable solid propellant rocket motors for the kick stage (auxiliary stage) of the Shuttle Tug, or Interim Upper Stage, are described, with details on features and test results of the ignition and quench (thrust termination) systems and procedures, fabrication of propellant and insulation, explosion hazards of propellants, and comparative data on present and future motor design. These rocket motor systems are designed for upper stage augmentation of launch vehicles and possible service in Shuttle-launched outer planet spacecraft.

  7. Simulation of liquid propellant rocket engine combustion instabilities

    Science.gov (United States)

    Ventrice, M. B.; Fang, J. C.; Purdy, K. R.

    1979-01-01

    A simulation technique for studying the high frequency combustion instabilities of liquid propellant rocket engines has been developed and used to investigate various aspects of instability phenomena. Of importance was investigation of the significance of the method of coupling the combustion and the gas dynamics of the system. Two coupling processes were studied: linear response of the combustion process to pressure fluctuations, and the nature of the resulting instabilities; and nonlinear response of the combustion process to velocity fluctuations, and the nature of the resulting instabilities. For the combustion model studied, nonlinear (velocity) coupling was found to more closely characterize liquid propellant instabilities.

  8. Development of hydrazinium nitroformate based solid propellants

    NARCIS (Netherlands)

    Schöyer, H.F.R.; Schnorhk, A.J.; Korting, P.A.O.G.; Lit, P.J. van; Mul, J.M.; Gadiot, G.; Meulenbrugge, J.J.

    1995-01-01

    The development of new high-performance propellant combinations requires the establishment of safety and handling characteristics and thermodynamic decomposition and explosive properties. This paper addresses the early development phases of a new composite solid propellant based on HNF as oxidizer

  9. Using PDV to Understand Damage in Rocket Motor Propellants

    Science.gov (United States)

    Tear, Gareth; Chapman, David; Ottley, Phillip; Proud, William; Gould, Peter; Cullis, Ian

    2017-06-01

    There is a continuing requirement to design and manufacture insensitive munition (IM) rocket motors for in-service use under a wide range of conditions, particularly due to shock initiation and detonation of damaged propellant spalled across the central bore of the rocket motor (XDT). High speed photography has been crucial in determining this behaviour, however attempts to model the dynamic behaviour are limited by the lack of precision particle and wave velocity data with which to validate against. In this work Photonic Doppler Velocimetery (PDV) has been combined with high speed video to give accurate point velocity and timing measurements of the rear surface of a propellant block impacted by a fragment travelling upto 1.4 km s-1. By combining traditional high speed video with PDV through a dichroic mirror, the point of velocity measurement within the debris cloud has been determined. This demonstrates a new capability to characterise the damage behaviour of a double base rocket motor propellant and hence validate the damage and fragmentation algorithms used in the numerical simulations.

  10. Solid Rocket Testing at AFRL (Briefing Charts)

    Science.gov (United States)

    2016-10-21

    19b. TELEPHONE NUMBER (Include area code) 10/21/2016 Briefing Charts 01 October 2016 - 31 October 2016 Solid Rocket Testing at AFRL Robert Antypas Air...Unclassified SAR 18 R. Antypas N/A Solid Rocket Testing at AFRL 21 Oct 2016 Robert Antypas AFRL/RQRO -Distribution A: Approved for Public Release...Distribution Unlimited. PA#16492 2 Agenda • Solid Rocket Motors • History of Sea Level Testing • Small Component Testing • Full-scale Testing • Altitude

  11. Scaling of Performance in Liquid Propellant Rocket Engine Combustion Devices

    Science.gov (United States)

    Hulka, James R.

    2008-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  12. Nitramine Composite Solid Propellant Modelling

    Science.gov (United States)

    1989-07-01

    flame occurs between the decomposition .products of the AP forming 02 and inert products. Thus: NH 3 + HCLO4 - Inert Products + 02 (26) Since the AP...Equilibrium Compositions, Thermodynamic Properties, and Performance Characteristics of Propellants Systems," NWC-TP-6037, April 1979. 152. Juhasz, A. A., "Round

  13. 40 CFR 61.43 - Emission testing-rocket firing or propellant disposal.

    Science.gov (United States)

    2010-07-01

    ... 40 Protection of Environment 8 2010-07-01 2010-07-01 false Emission testing-rocket firing or... Standard for Beryllium Rocket Motor Firing § 61.43 Emission testing—rocket firing or propellant disposal. (a) Ambient air concentrations shall be measured during and after firing of a rocket motor or...

  14. Computational Thermochemistry of Jet Fuels and Rocket Propellants

    Science.gov (United States)

    Crawford, T. Daniel

    2002-01-01

    The design of new high-energy density molecules as candidates for jet and rocket fuels is an important goal of modern chemical thermodynamics. The NASA Glenn Research Center is home to a database of thermodynamic data for over 2000 compounds related to this goal, in the form of least-squares fits of heat capacities, enthalpies, and entropies as functions of temperature over the range of 300 - 6000 K. The chemical equilibrium with applications (CEA) program written and maintained by researchers at NASA Glenn over the last fifty years, makes use of this database for modeling the performance of potential rocket propellants. During its long history, the NASA Glenn database has been developed based on experimental results and data published in the scientific literature such as the standard JANAF tables. The recent development of efficient computational techniques based on quantum chemical methods provides an alternative source of information for expansion of such databases. For example, it is now possible to model dissociation or combustion reactions of small molecules to high accuracy using techniques such as coupled cluster theory or density functional theory. Unfortunately, the current applicability of reliable computational models is limited to relatively small molecules containing only around a dozen (non-hydrogen) atoms. We propose to extend the applicability of coupled cluster theory- often referred to as the 'gold standard' of quantum chemical methods- to molecules containing 30-50 non-hydrogen atoms. The centerpiece of this work is the concept of local correlation, in which the description of the electron interactions- known as electron correlation effects- are reduced to only their most important localized components. Such an advance has the potential to greatly expand the current reach of computational thermochemistry and thus to have a significant impact on the theoretical study of jet and rocket propellants.

  15. Combustion of metal agglomerates in a solid rocket core flow

    Science.gov (United States)

    Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

    2013-12-01

    The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

  16. Development of small solid rocket boosters for the ILR-33 sounding rocket

    Science.gov (United States)

    Nowakowski, Pawel; Okninski, Adam; Pakosz, Michal; Cieslinski, Dawid; Bartkowiak, Bartosz; Wolanski, Piotr

    2017-09-01

    This paper gives an overview of the development of a 6000 Newton-class solid rocket motor for suborbital applications. The design configuration and results of interior ballistics calculations are given. The initial use of the motor as the main propulsion system of the H1 experimental in-flight test platform, within the Polish Small Sounding Rocket Program, is presented. Comparisons of theoretical and experimental performance are shown. Both on-ground and in-flight tests are discussed. A novel composite-case manufacturing technology, which enabled to reach high propellant mass fractions, was validated and significant cost-reductions were achieved. This paper focuses on the process of adapting the design for use as the booster stage of the ILR-33 sounding rocket, under development at the Institute of Aviation in Warsaw, Poland. Parallel use of two of the flight-proven rocket motors along with the main stage is planned. The process of adapting the rocket motor for booster application consists of stage integration, aerothermodynamics and reliability analyses. The separation mechanism and environmental impact are also discussed within this paper. Detailed performance analysis with focus on propellant grain geometry is provided. The evolution of the design since the first flights of the H1 rocket is covered and modifications of the manufacturing process are described. Issues of simultaneous ignition of two motors and their non-identical performance are discussed. Further applications and potential for future development are outlined. The presented results are based on the initial work done by the Rocketry Group of the Warsaw University of Technology Students' Space Association. The continuation of the Polish Small Sounding Rocket Program on a larger scale at the Institute of Aviation proves the value of the outcomes of the initial educational project.

  17. Holographic Investigation of Solid Propellant Combustion

    Science.gov (United States)

    1988-12-01

    IITLE (B=m* £Se-T) CaruatwICi, HOLOGRAPHIC INIL’ESTIGWI’ION OF SOLZ ’PROPELLANT COMBUSTION 12 PERSONAL AUTHOR(S) -Butler, Albert G. 13a TYPE OF REPORT...Speckle 19 ABSTRACT (Ccin1,rn WI ree~rse if IlCCenr~y anld ,dentify by blod number) An Investigcation into the behavior of aissmnied solid propellant...required to obtain a good quality hologram. Nuetral density filters, :3 placed in the scene beam for collimated type transmission holograms or in the

  18. Performance evaluation of commercial copper chromites as burning rate catalyst for solid propellants

    OpenAIRE

    Campos,Eunice Aparecida; Rita de Cássia L. Dutra; Rezende, Luis Cláudio; Diniz,Milton Faria; Nawa, Wilma Massae Dio; Iha,Koshun

    2010-01-01

    Abstract: Copper chromites are well known as burning rate catalysts for the combustion of composite solid propellants, used as a source of energy for rocket propulsion. The propellant burning rate depends upon the catalyst characteristics such as chemical composition and specific surface area. In this work, copper chromite samples from different suppliers were characterized by chemical analysis, FT-IR spectroscopy and by surface area measurement (BET). The samples were then evaluated as burni...

  19. Performance evaluation of commercial copper chromites as burning rate catalyst for solid propellants

    OpenAIRE

    Milton Faria Diniz; Eunice Aparecida Campos; Luis Cláudio Rezende; Rita de Cássia L. Dutra; Wilma Massae Dio Nawa; Koshun Iha

    2010-01-01

    Copper chromites are well known as burning rate catalysts for the combustion of composite solid propellants, used as a source of energy for rocket propulsion. The propellant burning rate depends upon the catalyst characteristics such as chemical composition and specific surface area. In this work, copper chromite samples from different suppliers were characterized by chemical analysis, FT-IR spectroscopy and by surface area measurement (BET). The samples were then evaluated as burning rate ca...

  20. Specific Impulses Losses in Solid Propellant Rockets

    Science.gov (United States)

    1974-12-17

    spectropyrometer enables the mean temperature of the particles to be found and infra- red emission to be measured; from this one deduces the gas temper...corresponds to phase equilibrium, we study the evolucion of velocity and temperature of the particles in this flow. We can thus calculate the inter-phase

  1. Unsteady Aerodynamic Investigation of the Propeller-Wing Interaction for a Rocket Launched Unmanned Air Vehicle

    Directory of Open Access Journals (Sweden)

    G. Q. Zhang

    2013-01-01

    Full Text Available The aerodynamic characteristics of propeller-wing interaction for the rocket launched UAV have been investigated numerically by means of sliding mesh technology. The corresponding forces and moments have been collected for axial wing placements ranging from 0.056 to 0.5D and varied rotating speeds. The slipstream generated by the rotating propeller has little effects on the lift characteristics of the whole UAV. The drag can be seen to remain unchanged as the wing's location moves progressively closer to the propeller until 0.056D away from the propeller, where a nearly 20% increase occurred sharply. The propeller position has a negligible effect on the overall thrust and torque of the propeller. The efficiency affected by the installation angle of the propeller blade has also been analyzed. Based on the pressure cloud and streamlines, the vortices generated by propeller, propeller-wing interaction, and wing tip have also been captured and analyzed.

  2. Combustion characteristics of a solid propellant with a charring binder

    Science.gov (United States)

    Udlock, D. E.; Strand, L. D.

    1973-01-01

    A brief investigation of the combustion characteristics of a solid propellant containing a binder which chars, as opposed to melting or volatizing, has been made. The burning rate of the propellant with the charring binder was significantly higher than similar propellants containing non-charring binders. High speed motion pictures of the burning propellant showed that the aluminum burned on the regressing surface, rather than a short distance from it as is typical with composite propellants.

  3. Solid State MEMS Thrusters Using Electrically Controlled Extinguishable Solid Propellant Project

    Data.gov (United States)

    National Aeronautics and Space Administration — ET Materials, LLC developed the first ever electrically controlled extinguishable solid propellant (ECESP). The original propellant developed under Air Force SBIR...

  4. Scale Effects on Solid Rocket Combustion Instability Behaviour

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2011-01-01

    Full Text Available The ability to understand and predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. A numerical model incorporating pertinent elements, such as a representative transient, frequency-dependent combustion response to pressure wave activity above the burning propellant surface, is applied to the investigation of scale effects (motor size, i.e., grain length and internal port diameter on influencing instability-related behaviour in a cylindrical-grain motor. The results of this investigation reveal that the motor’s size has a significant influence on transient pressure wave magnitude and structure, and on the appearance and magnitude of an associated base pressure rise.

  5. Scale effects on solid rocket combustion instability behaviour

    Energy Technology Data Exchange (ETDEWEB)

    Greatrix, D. R. [Ryerson University, Department of Aerospace Engineering, Toronto, Ontario (Canada)

    2011-07-01

    The ability to understand and predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. A numerical model incorporating pertinent elements, such as a representative transient, frequency-dependent combustion response to pressure wave activity above the burning propellant surface, is applied to the investigation of scale effects (motor size, i.e., grain length and internal port diameter) on influencing instability-related behaviour in a cylindrical-grain motor. The results of this investigation reveal that the motor's size has a significant influence on transient pressure wave magnitude and structure, and on the appearance and magnitude of an associated base pressure rise. (author)

  6. Solid Rocket Motor Insulation Testing

    Science.gov (United States)

    2005-07-13

    ablation test, analysis): $2K • Rapid testing of 5-6 samples per day (or more) •Nozzle materials can also be tested Propellant Test Insulation...Igniter Backing Nozzle Distribution A: Public Release. Distribution Unlimited 4 Distribution A: Public Release. Distribution unlimited Nanofilled EPDM as...Improved Ablative Potential Pathway to Increase Performance M i c r o f i b e r s SOTA EPDM Rubber NANO EPDM Rubber ... . .. .burn carbon char

  7. Distinctive features of the intrachamber instability of combustion in liquid-propellant rocket engines

    Science.gov (United States)

    Gotsulenko, V. V.

    2008-09-01

    Self-oscillations and certain of their regularities determined by solution of a degenerate system of differential equations that is used in considering combustion instability in combustion chambers of liquid-propellant rocket engines are modeled mathematically.

  8. Studies on Pre-Ignition Chamber Dynamics of Solid Rockets with Different Port Geometries

    OpenAIRE

    S. Vivek; Sharad Sharan; R. Arvind; D. V. Praveen; J. Vigneshwar; S. Ajith; V. R. Sanal Kumar

    2015-01-01

    In this paper numerical studies have been carried out to examine the pre-ignition flow features of high-performance solid propellant rocket motors with two different port geometries but with same propellant loading density. Numerical computations have been carried out using a validated 3D, unsteady, 2nd-order implicit, SST k- ω turbulence model. In the numerical study, a fully implicit finite volume scheme of the compressible, Reynolds-Averaged, Navier- Stokes equations i...

  9. Mixing and combustion enhancement of Turbocharged Solid Propellant Ramjet

    Science.gov (United States)

    Liu, Shichang; Li, Jiang; Zhu, Gen; Wang, Wei; Liu, Yang

    2018-02-01

    Turbocharged Solid Propellant Ramjet is a new concept engine that combines the advantages of both solid rocket ramjet and Air Turbo Rocket, with a wide operation envelope and high performance. There are three streams of the air, turbine-driving gas and augment gas to mix and combust in the afterburner, and the coaxial intake mode of the afterburner is disadvantageous to the mixing and combustion. Therefore, it is necessary to carry out mixing and combustion enhancement research. In this study, the numerical model of Turbocharged Solid Propellant Ramjet three-dimensional combustion flow field is established, and the numerical simulation of the mixing and combustion enhancement scheme is conducted from the aspects of head region intake mode to injection method in afterburner. The results show that by driving the compressed air to deflect inward and the turbine-driving gas to maintain strong rotation, radial and tangential momentum exchange of the two streams can be enhanced, thereby improving the efficiency of mixing and combustion in the afterburner. The method of injecting augment gas in the transverse direction and making sure the injection location is as close as possible to the head region is beneficial to improve the combustion efficiency. The outer combustion flow field of the afterburner is an oxidizer-rich environment, while the inner is a fuel-rich environment. To improve the efficiency of mixing and combustion, it is necessary to control the injection velocity of the augment gas to keep it in the oxygen-rich zone of the outer region. The numerical simulation for different flight conditions shows that the optimal mixing and combustion enhancement scheme can obtain high combustion efficiency and have excellent applicability in a wide working range.

  10. Advanced Solid Rocket Motor case design status

    Science.gov (United States)

    Palmer, G. L.; Cash, S. F.; Beck, J. P.

    1993-06-01

    The Advanced Solid Rocket Motor (ASRM) case design aimed at achieving a safer and more reliable solid rocket motor for the Space Shuttle system is considered. The ASRM case has a 150.0 inch diameter, three equal length segment, and 9Ni-4CO-0.3C steel alloy. The major design features include bolted casebolted case joints which close during pressurization, plasma arc welded factory joints, integral stiffener for splash down and recovery, and integral External Tank attachment rings. Each mechanical joint has redundant and verifiable o-ring seals.

  11. Method and device for protecting objects against Rocket Propelled Grenades (RPGs)

    NARCIS (Netherlands)

    Wentzel, C.M.

    2008-01-01

    A device and a method for protecting objects against rocket-propelled grenades having a hollow nose cone includes a netting of knotted and coated superstrong fibers disposed in front of the object, in such a manner that the nose cone of a rocket caught in the netting will penetrate one of the meshes

  12. Feasibility Study and Demonstration of an Aluminum and Ice Solid Propellant

    Directory of Open Access Journals (Sweden)

    Timothee L. Pourpoint

    2012-01-01

    Full Text Available Aluminum-water reactions have been proposed and studied for several decades for underwater propulsion systems and applications requiring hydrogen generation. Aluminum and water have also been proposed as a frozen propellant, and there have been proposals for other refrigerated propellants that could be mixed, frozen in situ, and used as solid propellants. However, little work has been done to determine the feasibility of these concepts. With the recent availability of nanoscale aluminum, a simple binary formulation with water is now feasible. Nanosized aluminum has a lower ignition temperature than micron-sized aluminum particles, partly due to its high surface area, and burning times are much faster than micron aluminum. Frozen nanoscale aluminum and water mixtures are stable, as well as insensitive to electrostatic discharge, impact, and shock. Here we report a study of the feasibility of an nAl-ice propellant in small-scale rocket experiments. The focus here is not to develop an optimized propellant; however improved formulations are possible. Several static motor experiments have been conducted, including using a flight-weight casing. The flight weight casing was used in the first sounding rocket test of an aluminum-ice propellant, establishing a proof of concept for simple propellant mixtures making use of nanoscale particles.

  13. Introduction to rocket science and engineering

    CERN Document Server

    Taylor, Travis S

    2009-01-01

    What Are Rockets? The History of RocketsRockets of the Modern EraRocket Anatomy and NomenclatureWhy Are Rockets Needed? Missions and PayloadsTrajectoriesOrbitsOrbit Changes and ManeuversBallistic Missile TrajectoriesHow Do Rockets Work? ThrustSpecific ImpulseWeight Flow RateTsiolkovsky's Rocket EquationStagingRocket Dynamics, Guidance, and ControlHow Do Rocket Engines Work? The Basic Rocket EngineThermodynamic Expansion and the Rocket NozzleExit VelocityRocket Engine Area Ratio and LengthsRocket Engine Design ExampleAre All Rockets the Same? Solid Rocket EnginesLiquid Propellant Rocket Engines

  14. Unique thermocouple to measure the temperatures of squibs, igniters, propellants, and rocket nozzles

    Science.gov (United States)

    Nanigian, Jacob; Nanigian, Dan

    2006-05-01

    The temperatures produced by the various components in the propulsion system of rockets and missiles determine the performance of the rocket. Since these temperatures occur very rapidly and under extreme conditions, standard thermocouples fail before any meaningful temperatures are measured. This paper describes the features of a special family of high performance thermocouples, which can measure these transient temperatures with millisecond response times and under the most severe conditions of erosion. Examples of igniter, propellant and rocket nozzle temperatures are included in this paper. Also included is heat flux measurements made by these sensors in rocket applications.

  15. Evolution of solid rocket booster component testing

    Science.gov (United States)

    Lessey, Joseph A.

    1989-01-01

    The evolution of one of the new generation of test sets developed for the Solid Rocket Booster of the U.S. Space Transportation System. Requirements leading to factory checkout of the test set are explained, including the evolution from manual to semiautomated toward fully automated status. Individual improvements in the built-in test equipment, self-calibration, and software flexibility are addressed, and the insertion of fault detection to improve reliability is discussed.

  16. Performance evaluation of commercial copper chromites as burning rate catalyst for solid propellants

    Directory of Open Access Journals (Sweden)

    Milton Faria Diniz

    2010-09-01

    Full Text Available Copper chromites are well known as burning rate catalysts for the combustion of composite solid propellants, used as a source of energy for rocket propulsion. The propellant burning rate depends upon the catalyst characteristics such as chemical composition and specific surface area. In this work, copper chromite samples from different suppliers were characterized by chemical analysis, FT-IR spectroscopy and by surface area measurement (BET. The samples were then evaluated as burning rate catalyst in a typical composite propellant formulation based on HTPB binder, ammonium perchlorate and aluminum. The obtained surface area values are very close to those informed by the catalyst suppliers. The propellant processing as well as its mechanical properties were not substantially affected by the type of catalyst. Some copper chromite catalysts caused an increase in the propellant burning rate in comparison to the iron oxide catalyst. The results show that in addition to the surface area, other parameters like chemical composition, crystalline structure and the presence of impurities might be affecting the catalyst performance. All evaluated copper chromite samples may be used as burning rate catalyst in composite solid propellant formulations, with slight advantages for the SX14, Cu-0202P and Cu-1800P samples, which led to the highest burning rate propellants.

  17. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Under a Phase 1 effort, IES successfully developed and demonstrated a spark ignition concept where propellant flow drives a very simple fluid mechanical oscillator...

  18. Firing Tests of Thrust Controllable Solid Rocket Motors Part 1:300-mm-Diameter End-Burning Grain Motors

    OpenAIRE

    GODAI, Tomifumi; SHIMIZU, Morio; Ito, Katsuya; Tanemura, Toshiharu; Fujiwara, Tsutomu; KUSAKA, Kazuo; Kisara, Katsuto; Sato, Masahiro; Takahashi, Mamoru; Izumikawa, Muneo; 五代, 富文; 清水, 盛生; 伊藤, 克弥; 種村, 利春; 藤原, 勉

    1980-01-01

    300-mm-diameter thrust controllable solid rocket motors were successfully fired under atmospheric conditions. Thrust magnitude control was attained by changing the flow rate of the hydrogen peroxide decomposition gas injected into the solid rocket motor. Incomplete combustion termination that occurred earlier than expected was due to the excessive heat flux irradiated from the hot surfaces of the ablative insulation liner towards the extinguished propellant surface. More effective liner mater...

  19. Liquid-propellant rocket engine testing at Arnold Engineering Development Center

    Science.gov (United States)

    Dehoff, Bryan; Tucker, Edgar K.; McAmis, Rob W.

    A continuing need exists for facilities to test both storable and cryogenic liquid-propellant rocket engines and stages at simulated altitude as part of a resposible acquisition risk reduction program. Storable propellant Intercontinental Ballistic Missile (ICBM) Post Boost Vehicles (PBV) require simulated altitude testing as part of the Aging and Surveillance programs designed to ensure an effective and reliable missile system. Likewise, simulated altitude testing is necessary acquisition risk reduction for advanced cryogenic rocket engines and stages that are being upgraded or developed to satisfy a variety of defense and commercial payload requirements. A review of liquid rocket test facilities at Arnold Engineering Development Center (AEDC) is presented. The facility capabilities used in support of acquisition risk reduction are described, as are new facility capabilities recently completed or funded. Furthermore, a description of the technology applications available at AEDC in support of liquid rocket diagnostics, analysis, and evaluation techniques is presented.

  20. Crusader solid propellant best technical approach

    Energy Technology Data Exchange (ETDEWEB)

    Graves, V. [Oak Ridge National Lab., TN (United States); Bader, G. [Fire Support Armament Center, Picatinny Arsenal, NJ (United States); Dolecki, M. [Tank-Automotive Research, Development, and Engineering Center, Picatinny, NJ (United States); Krupski, S. [Benet Weapons Lab., Watervliet Arsenal, NY (United States); Zangrando, R. [Close Combat Armament Center, Picatinny Arsenal, NJ (United States)

    1995-12-01

    The goal of the Solid Propellant Resupply Team is to develop Crusader system concepts capable of automatically handling 155mm projectiles and Modular Artillery Charges (MACs) based on system requirements. The system encompasses all aspects of handling from initial input into a resupply vehicle (RSV) to the final loading into the breech of the self-propelled howitzer (SPH). The team, comprised of persons from military and other government organizations, developed concepts for the overall vehicles as well as their interior handling components. An intermediate review was conducted on those components, and revised concepts were completed in May 1995. A concept evaluation was conducted on the finalized concepts, from both a systems level and a component level. The team`s Best Technical Approach (BTA) concept was selected from that evaluation. Both vehicles in the BTA have a front-engine configuration with the crew situated behind the engine-low in the vehicles. The SPH concept utilizes an automated reload port at the rear of the vehicle, centered high. The RSV transfer boom will dock with this port to allow automated ammunition transfer. The SPH rearm system utilizes fully redundant dual loaders. Active magazines are used for both projectiles and MACs. The SPH also uses a nonconventional tilted ring turret configuration to maximize the available interior volume in the vehicle. This configuration can be rearmed at any elevation angle but only at 0{degree} azimuth. The RSV configuration is similar to that of the SPH. The RSV utilizes passive storage racks with a pick-and-place manipulator for handling the projectiles and active magazines for the MACs. A telescoping transfer boom extends out the front of the vehicle over the crew and engine.

  1. Study of Liquid Breakup Process in Solid Rocket Motors

    Science.gov (United States)

    2014-01-01

    Fellowship Program (SFFP). V REFERENCES                                                                  1 Xiao, Y.M., R.S. Amano, " Aluminized ...Motors," Journal of Propulsion and Power, Vol.11, pp. 10-23, 1995. 8 Hess, E., Chen, K., Acosta, P., Brent, D., Fendell, F.; "Effect of Aluminized ...No. 9, pp. 1763-1770, September 2003. 11 Bandera, A., Maggi, F., and Deluca, L.T., "Agglomeration of Aluminized Solid Rocket Propellants," in 45th

  2. Studies on composite solid propellant with tri-modal ammonium perchlorate containing an ultrafine fraction

    Directory of Open Access Journals (Sweden)

    K.V. Suresh Babu

    2017-08-01

    composite solid propellant is prepared by using burn rate modifiers Copper chromite and Iron oxide. Addition of Copper chromite and Iron oxide has enhanced the burn rate of tri-modal AP based composite solid propellant. The catalytic propensity of copper chromite is higher than that of iron oxide. The pressure exponent increased with the catalyst concentration and the values obtained are compatible for solid rocket motor applications.

  3. A two-phase restricted equilibrium model for combustion of metalized solid propellants

    Science.gov (United States)

    Sabnis, J. S.; Dejong, F. J.; Gibeling, H. J.

    1992-01-01

    An Eulerian-Lagrangian two-phase approach was adopted to model the multi-phase reacting internal flow in a solid rocket with a metalized propellant. An Eulerian description was used to analyze the motion of the continuous phase which includes the gas as well as the small (micron-sized) particulates, while a Lagrangian description is used for the analysis of the discrete phase which consists of the larger particulates in the motor chamber. The particulates consist of Al and Al2O3 such that the particulate composition is 100 percent Al at injection from the propellant surface with Al2O3 fraction increasing due to combustion along the particle trajectory. An empirical model is used to compute the combustion rate for agglomerates while the continuous phase chemistry is treated using chemical equilibrium. The computer code was used to simulate the reacting flow in a solid rocket motor with an AP/HTPB/Al propellant. The computed results show the existence of an extended combustion zone in the chamber rather than a thin reaction region. The presence of the extended combustion zone results in the chamber flow field and chemical being far from isothermal (as would be predicted by a surface combustion assumption). The temperature in the chamber increases from about 2600 K at the propellant surface to about 3350 K in the core. Similarly the chemical composition and the density of the propellant gas also show spatially non-uniform distribution in the chamber. The analysis developed under the present effort provides a more sophisticated tool for solid rocket internal flow predictions than is presently available, and can be useful in studying apparent anomalies and improving the simple correlations currently in use. The code can be used in the analysis of combustion efficiency, thermal load in the internal insulation, plume radiation, etc.

  4. Regarding the perturbed operating process of DB propellant rocket motor at extreme initial grain temperatures

    Directory of Open Access Journals (Sweden)

    Ioan ION

    2012-03-01

    Full Text Available Despite many decades of study, the combustion instability of several DB propellants is still of particular concern, especially at extreme grain temperature conditions of rocket motor operating. The purpose of the first part of the paper is to give an overview of our main experimental results on combustion instabilities and pressure oscillations in DB propellant segmented grain rocket motors (SPRM-01, large L/D ratio, working at extreme initial grain temperatures. Thus, we recorded some particular pressure-time traces with significant perturbed pressure signal that was FFT analysed. An updated mathematical model incorporating transient frequency-dependent combustion response, in conjunction with pressure-dependent burning, is applied to investigate and predict the DB propellant combustion instability phenomenon. The susceptibility of the tested motor SPRM-01 with DB propellant to get a perturbed working and to go unstable with pressure was evidenced and this risk has to be evaluated. In the last part of our paper we evaluated the influence of recorded perturbed thrust on the rocket behaviour on the trajectory. The study revealed that at firing-table initial conditions, this kind of perturbed motor operating may not lead to an unstable rocket flight, but the ballistic parameters would be influenced in an unacceptable manner.

  5. Combustion stability with baffles, absorbers and velocity sensitive combustion. [liquid propellant rocket combustors

    Science.gov (United States)

    Mitchell, C. E.

    1980-01-01

    Analytical and computational techniques were developed to predict the stability behavior of liquid propellant rocket combustors using damping devices such as acoustic liners, slot absorbers, and injector face baffles. Models were developed to determine the frequency and decay rate of combustor oscillations, the spatial and temporal pressure waveforms, and the stability limits in terms of combustion response model parameters.

  6. Radiation from advanced solid rocket motor plumes

    Science.gov (United States)

    Farmer, Richard C.; Smith, Sheldon D.; Myruski, Brian L.

    1994-12-01

    The overall objective of this study was to develop an understanding of solid rocket motor (SRM) plumes in sufficient detail to accurately explain the majority of plume radiation test data. Improved flowfield and radiation analysis codes were developed to accurately and efficiently account for all the factors which effect radiation heating from rocket plumes. These codes were verified by comparing predicted plume behavior with measured NASA/MSFC ASRM test data. Upon conducting a thorough review of the current state-of-the-art of SRM plume flowfield and radiation prediction methodology and the pertinent data base, the following analyses were developed for future design use. The NOZZRAD code was developed for preliminary base heating design and Al2O3 particle optical property data evaluation using a generalized two-flux solution to the radiative transfer equation. The IDARAD code was developed for rapid evaluation of plume radiation effects using the spherical harmonics method of differential approximation to the radiative transfer equation. The FDNS CFD code with fully coupled Euler-Lagrange particle tracking was validated by comparison to predictions made with the industry standard RAMP code for SRM nozzle flowfield analysis. The FDNS code provides the ability to analyze not only rocket nozzle flow, but also axisymmetric and three-dimensional plume flowfields with state-of-the-art CFD methodology. Procedures for conducting meaningful thermo-vision camera studies were developed.

  7. Large Solid Rocket Motor Safety Analyses: Thermal Effects Issues

    Science.gov (United States)

    2010-07-01

    source (industrial site…), hazard flux (propagation, toxic species dispersion, fire and explosion) and targets submitted to damage (personnel, material...and two propellant compositions (20% and 4% of Aluminium in mass) are tested within this first setup. The solid propellants are produced by the SME...and it is 1.10 mm/s ±0.04 mm/s with the 4% Aluminium one. In other words, higher aluminium rate in mass is, faster the solid propellant burns

  8. Applied algorithm in the liner inspection of solid rocket motors

    Science.gov (United States)

    Hoffmann, Luiz Felipe Simões; Bizarria, Francisco Carlos Parquet; Bizarria, José Walter Parquet

    2018-03-01

    In rocket motors, the bonding between the solid propellant and thermal insulation is accomplished by a thin adhesive layer, known as liner. The liner application method involves a complex sequence of tasks, which includes in its final stage, the surface integrity inspection. Nowadays in Brazil, an expert carries out a thorough visual inspection to detect defects on the liner surface that may compromise the propellant interface bonding. Therefore, this paper proposes an algorithm that uses the photometric stereo technique and the K-nearest neighbor (KNN) classifier to assist the expert in the surface inspection. Photometric stereo allows the surface information recovery of the test images, while the KNN method enables image pixels classification into two classes: non-defect and defect. Tests performed on a computer vision based prototype validate the algorithm. The positive results suggest that the algorithm is feasible and when implemented in a real scenario, will be able to help the expert in detecting defective areas on the liner surface.

  9. Process and quality verification controls for solid propellant manufacturing

    Science.gov (United States)

    Rogers, C. J.

    1983-01-01

    It is pointed out that in-process tests to verify quality and detect discrepant propellant which could compromise motor performance are essential elements of the solid composite propellant manufacturing process. The successful performance of the 260SL-1 and 260SL-2 motors aptly verified the controls used for manufacturing the propellant. The present investigation is concerned with the selected control parameters, and their relationships to composition and final propellant properties. Control performance is evaluated by comparison with processing data experienced in the manufacture of the propellant for the 260SL-1 motor. It is found that the in-process quality verification controls utilized in the propellant manufacturing process for the 260-in. diameter motor contributed significantly to the confidence of successful and predictable motor performance.

  10. Experimental validation of solid rocket motor damping models

    Science.gov (United States)

    Riso, Cristina; Fransen, Sebastiaan; Mastroddi, Franco; Coppotelli, Giuliano; Trequattrini, Francesco; De Vivo, Alessio

    2017-12-01

    In design and certification of spacecraft, payload/launcher coupled load analyses are performed to simulate the satellite dynamic environment. To obtain accurate predictions, the system damping properties must be properly taken into account in the finite element model used for coupled load analysis. This is typically done using a structural damping characterization in the frequency domain, which is not applicable in the time domain. Therefore, the structural damping matrix of the system must be converted into an equivalent viscous damping matrix when a transient coupled load analysis is performed. This paper focuses on the validation of equivalent viscous damping methods for dynamically condensed finite element models via correlation with experimental data for a realistic structure representative of a slender launch vehicle with solid rocket motors. A second scope of the paper is to investigate how to conveniently choose a single combination of Young's modulus and structural damping coefficient—complex Young's modulus—to approximate the viscoelastic behavior of a solid propellant material in the frequency band of interest for coupled load analysis. A scaled-down test article inspired to the Z9-ignition Vega launcher configuration is designed, manufactured, and experimentally tested to obtain data for validation of the equivalent viscous damping methods. The Z9-like component of the test article is filled with a viscoelastic material representative of the Z9 solid propellant that is also preliminarily tested to investigate the dependency of the complex Young's modulus on the excitation frequency and provide data for the test article finite element model. Experimental results from seismic and shock tests performed on the test configuration are correlated with numerical results from frequency and time domain analyses carried out on its dynamically condensed finite element model to assess the applicability of different equivalent viscous damping methods to describe

  11. Detailed modal testing of a solid rocket motor using a portable test system

    Science.gov (United States)

    Glozman, Vladimir; Brillhart, Ralph D.

    1990-01-01

    Modern analytical techniques have expended the ability to evaluate solid rocket motors used in launch vehicles. As more detailed models of solid rocket motors were developed, testing methods were required to verify the models. Experimental modal analysis (modal testing) of space structures and launch vehicles has been a requirement for model validation for many years. However, previous testing of solid rocket motors has not typically involved dynamic modal testing of full scale motors for verification of solid propellant or system assembly properties. Innovative approaches to the testing of solid rocket motors were developed and modal testing of a full scale, two segment Titan 34D Solid Rocket Motor (SRM) was performed to validate detailed computer modeling. Special modifications were made to convert an existing facility into a temporary modal test facility which would accommodate the test article. The assembly of conventional data acquisition equipment into a multiple channel count portable system has made modal testing in the field feasible. Special purpose hydraulic exciters were configured to apply the dynamic driving forces required. All instrumentation and data collection equipment were installed at the test site for the duration of the test program and removed upon completion. Conversion of an existing test facility into a temporary modal test facility, and use of a multiple channel count portable test data acquisition system allowed all test objectives to be met and resulted in validation of the computer model in a minimum time.

  12. Numerical Evaluation of the Use of Aluminum Particles for Enhancing Solid Rocket Motor Combustion Stability

    OpenAIRE

    David Greatrix

    2015-01-01

    The ability to predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave ...

  13. Assessment of tbe Performance of Ablative Insulators Under Realistic Solid Rocket Motor Operating Conditions (a Doctoral Dissertation)

    Science.gov (United States)

    Martin, Heath Thomas

    2013-01-01

    Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

  14. Attenuation studies of booster-rocket propellants and their simulants

    Energy Technology Data Exchange (ETDEWEB)

    Weirick, L.J.

    1990-08-01

    A series of impact experiments on a composite propellant, an energetic propellant, and their simulants was recently completed using a light-gas gun. Previous experiments were done to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall damage at various impact velocities. The present studies measured the attenuation of shock waves in these materials, completing the shock characterization needed for material modeling. An initial impulse of 2.0 GPa magnitude and {approximately}0.6 {mu}s duration was imposed upon samples of various thicknesses. VISAR was used to measure the free-surface velocity at the back of the samples; these data were used to generate a curve of shock-wave attenuation versus sample thickness for each material. Results showed that all four materials attenuated the shock wave very similarly. Material thicknesses of 3.0, 7.62, 12.7, and 19.0 mm attenuated the shock wave {approximately}16%, 33%, 50%, and 66% respectively. 14 refs., 12 figs., 4 tabs.

  15. Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

    Science.gov (United States)

    Hulka, James R.; Jones, Gregg W.

    2010-01-01

    Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using

  16. Modeling of cooling channel flow in liquid-propellant rocket engines

    OpenAIRE

    Pizzarelli, Marco

    2008-01-01

    Ever since the development of liquid rocket engine, there has been a need to predict the peak heat flux that affects the engine material and thus to control the wall thermal behavior of rocket engine. To prevent thermal failure, the engine is generally cooled by means of a coolant that flows in passages that line the hottest part of the engine (i.e., combustion chamber and nozzle wall). This is the fluid-cooling system. If the coolant is one of the propellants, once it passes through th...

  17. Combustion instability analysis for liquid propellant rocket engines

    Science.gov (United States)

    Kim, Y. M.; Chen, C. P.; Ziebarth, J. P.

    1992-01-01

    The multi-dimensional numerical model has been developed to analyze the nonlinear combustion instabilities in liquid-fueled engines. The present pressure-based approach can handle the implicit pressure-velocity coupling in a non-iterative way. The additional scalar conservation equations for the chemical species, the energy, and the turbulent transport quantities can be handled by the same predictor-corrector sequences. This method is time-accurate and it can be applicable to the all-speed, transient, multi-phase, and reacting flows. Special emphasis is given to the acoustic/vaporization interaction which may act as the crucial rate-controlling mechanism in the liquid-fueled rocket engines. The subcritical vaporization is modeled to account for the effects of variable thermophysical properties, non-unitary Lewis number in the gas-film, the Stefan flow effect, and the effect of transient liquid heating. The test cases include the one-dimenisonal fast transient non-reacting and reacting flows, and the multi-dimensional combustion instabilities encountered in the liquid-fueled rocket thrust chamber. The present numerical model successfully demonstrated the capability to simulate the fast transient spray-combusting flows in terms of the limiting-cycle amplitude phenomena, correspondence between combustion and acoustics, and the steep-fronted wave and flame propagation. The investigated parameters include the spray initial conditions, air-fuel mixture ratios, and the engine geometry. Stable and unstable operating conditions are found for the liquid-fueled combustors. Under certain conditions, the limiting cycle behavior of the combusting flowfields is obtained. The numerical results indicate that the spray vaporization processes play an important role in releasing thermal energy and driving the combustion instability.

  18. Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes

    Directory of Open Access Journals (Sweden)

    Jaime Rubio Hervas

    2012-01-01

    Full Text Available The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

  19. Performance and Stability Analyses of Rocket Combustion Devices Using Liquid Oxygen/Liquid Methane Propellants

    Science.gov (United States)

    Hulka, James R.; Jones, G. W.

    2010-01-01

    Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented programs with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations on these programs. This paper summarizes these analyses. Test and analysis results of impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Several cases with gaseous methane are included for reference. Several different thrust chamber configurations have been modeled, including thrust chambers with multi-element like-on-like and swirl coax element injectors tested at NASA MSFC, and a unielement chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods.

  20. Solid Rocket Motor Design Using Hybrid Optimization

    Directory of Open Access Journals (Sweden)

    Kevin Albarado

    2012-01-01

    Full Text Available A particle swarm/pattern search hybrid optimizer was used to drive a solid rocket motor modeling code to an optimal solution. The solid motor code models tapered motor geometries using analytical burn back methods by slicing the grain into thin sections along the axial direction. Grains with circular perforated stars, wagon wheels, and dog bones can be considered and multiple tapered sections can be constructed. The hybrid approach to optimization is capable of exploring large areas of the solution space through particle swarming, but is also able to climb “hills” of optimality through gradient based pattern searching. A preliminary method for designing tapered internal geometry as well as tapered outer mold-line geometry is presented. A total of four optimization cases were performed. The first two case studies examines designing motors to match a given regressive-progressive-regressive burn profile. The third case study studies designing a neutrally burning right circular perforated grain (utilizing inner and external geometry tapering. The final case study studies designing a linearly regressive burning profile for right circular perforated (tapered grains.

  1. Holographic Investigation of Solid Propellant Particulates.

    Science.gov (United States)

    1981-12-01

    used at the Naval Postgraduate School in an attempt to obtain this type of data. They are: 1. High speed cinematography of burning propellant strands in...techniques vice conventional photography. Utilization of the holographic procedure results in a film plate which has recorded on it both the phase and...AGFA-GEVAERT 8E75 HD film plate is mounted on a kinematic plate holder near the focal plane of a pair of plano convex lenses. This device serves to

  2. The Effect of Solid Constituent Particle Size Distributions on TP-H1148 Propellant Slag

    Science.gov (United States)

    May, Douglas H.; Miles, William L.; Taylor, David S.; Rackham, Jon L.

    1997-01-01

    Special aluminum and ammonium perchlorate (AP) particle size distributions were prepared for a matrix of five-inch diameter, center-perforated (CP) motor tests to measure the aluminum oxide slag response in Space Shuttle Reusable Solid Rocket Motor (RSRM) propellant. Previous tests of TP-H1148 propellant in five-inch CP spin motors have shown a correlation between aluminum particle size and generated slag. The motors for this study were cast from thirteen five-gallon propellant mixes which used five particle size levels of aluminum powder, five of unground AP and three of ground AP. Aluminum had the greatest effect on slag formation, the more coarse fractions causing greater slag quantities and larger slag particles. Unground AP had about half the effect of aluminum with the coarser fractions again producing more and larger sized slag particles. The variation in ground AP did not have a significant effect on slag formation. Quench bomb tests showed the same trends as the spin motors, that is, larger aluminum and AP particle size distributions generated larger slag particles leaving the propellant surface. Cured propellant mechanical properties were also impacted by particle size variation.

  3. Investigation on liquid helium pressurization process using a heater in a liquid propellant rocket

    Energy Technology Data Exchange (ETDEWEB)

    Sehwan In; Sangkwon Jeong; Hongseong Kim [Korea Advanced Inst. of Science and Technology, Dept. of Mechanical Engineering, Daejeon (Korea)

    2004-08-01

    This paper describes a novel concept of the liquid helium pressurization system utilizing a heater in a liquid propellant rocket. This method makes the pressurization system simple and reliable, because electrical energy can be easily controlled to maintain the expulsion pressure in the liquid helium tank. The experimental results and the thermodynamic analysis are presented in this paper. The thermodynamic analysis considers two control volumes each with a uniform temperature, with the pressure in them being equal. The experimental result shows that this pressurization method satisfies the requirement for the pressurization system of the rocket. The thermodynamic analysis is verified through the comparison with the experimental result. The electrical power requirement in a practical rocket is also discussed. (Author)

  4. Aluminum/hydrocarbon gel propellants: An experimental and theoretical investigation of secondary atomization and predicted rocket engine performance

    Science.gov (United States)

    Mueller, Donn Christopher

    1997-12-01

    Experimental and theoretical investigations of aluminum/hydrocarbon gel propellant secondary atomization and its potential effects on rocket engine performance were conducted. In the experimental efforts, a dilute, polydisperse, gel droplet spray was injected into the postflame region of a burner and droplet size distributions was measured as a function of position above the burner using a laser-based sizing/velocimetry technique. The sizing/velocimetry technique was developed to measure droplets in the 10-125 mum size range and avoids size-biased detection through the use of a uniformly illuminated probe volume. The technique was used to determine particle size distributions and velocities at various axial locations above the burner for JP-10, and 50 and 60 wt% aluminum gels. Droplet shell formation models were applied to aluminum/hydrocarbon gels to examine particle size and mass loading effects on the minimum droplet diameter that will permit secondary atomization. This diameter was predicted to be 38.1 and 34.7 mum for the 50 and 60 wt% gels, which is somewhat greater than the experimentally measured 30 and 25 mum diameters. In the theoretical efforts, three models were developed and an existing rocket code was exercised to gain insights into secondary atomization. The first model was designed to predict gel droplet properties and shell stresses after rigid shell formation, while the second, a one-dimensional gel spray combustion model was created to quantify the secondary atomization process. Experimental and numerical comparisons verify that secondary atomization occurs in 10-125 mum diameter particles although an exact model could not be derived. The third model, a one-dimensional gel-fueled rocket combustion chamber, was developed to evaluate secondary atomization effects on various engine performance parameters. Results show that only modest secondary atomization may be required to reduce propellant burnout distance and radiation losses. A solid propellant

  5. 14 CFR 420.65 - Handling of solid propellants.

    Science.gov (United States)

    2010-01-01

    ... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Handling of solid propellants. 420.65 Section 420.65 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL AVIATION ADMINISTRATION... from the closest debris or explosive hazard source in an explosive hazard facility. ...

  6. Developments on HNF based high performance and green solid propellants

    NARCIS (Netherlands)

    Keizers, H.L.J.; Heijden, A.E.D.M. van der; Vliet, L.D. van; Welland-Veltmans, W.H.M.; Ciucci, A.

    2001-01-01

    Worldwide developments are ongoing to develop new and more energetic composite solid propellant formulations for space transportation and military applications. Since the 90's, the use of HNF as a new high performance oxidiser is being reinvestigated. Within European development programmes,

  7. C/C-SiC Composites for Nozzle of Solid Propellant Ramjet

    Directory of Open Access Journals (Sweden)

    WANG Lingling

    2017-01-01

    Full Text Available Carbon fiber reinforced carbon and silicon carbide matrix composites for nozzle inner of solid propellant ramjet were prepared by using the hybrid process of "chemical vapor infiltration + precursor impregnation pyrolysis (CVI+PIP". The microstructure, flexural and anti-ablation properties of the C/C-SiC composites and hydraulic test and rocket motor hot firing test for nozzle inner of solid propellant ramjet were comprehensively investigated. The results show that when the flexural strength of the composite reachs 197 MPa, the fracture damage behavior of the composites presents typical toughness mode.Also the composites has excellent anti-ablative property, i.e., linear ablation rate is only 0.0063 mm·s-1 after 200 s ablation. The C/C-SiC component have excellent integral bearing performance with the hydraulic bursting pressure of 6.5 MPa, and the high temperature combination property of the C/C-SiC composite nozzle inner is verified through motor hot firing of solid propellant ramjet.

  8. Low-Cost High-Performance Non-Toxic Self-Pressurizing Storable Liquid Bi-Propellant Pressure-Fed Rocket Engine Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Exquadrum proposes a high-performance liquid bi-propellant rocket engine that uses propellants that are non-toxic, self-pressurizing, and low cost. The proposed...

  9. Study of Liquid Breakup Process in Solid Rocket Motor Nozzle

    Science.gov (United States)

    2016-02-16

    Journal Article 3. DATES COVERED (From - To) 16 June 2015 – 31 July 2015 4. TITLE AND SUBTITLE Study of Liquid Breakup Process in Solid Rocket Motor... Rockets , Volume 52 Issue 4 July 2015 (AIAA 2015-1605) PA Case Number: #15348; Clearance Date: 6/26/2015 ©2015 AIAA The U.S. Government is joint author...of the work and has the right to use, modify, reproduce, release, perform, display, or 14. ABSTRACT In a solid rocket motor (SRM), when the

  10. The measurement, modelling and prediction of traction forces in a rocket propellant

    Science.gov (United States)

    Tevaarwerk, J. L.; Woods, C.

    1990-01-01

    Traction tests were performed on RPI, a common kerosene-based rocket propellant. Traction data on this fluid are required for purposes of turbopump bearing design, using such codes as SHABERTH. To obtain the traction data, an existing twin disk machine was used, operating under the side slip mode and using elliptical contacts. The resulting traction curves were reduced to fundamental fluid property parameters using the Johnson and Tevaarwerk traction model. Theoretical traction predictions were performed by back substitution of the fundamental properties into the traction model. Comparison of the predicted with the measured curves gives a high degree of confidence in the correctness of the traction model.

  11. Control-oriented modeling of combustion and flow processes in liquid propellant rocket engines

    Science.gov (United States)

    Bentsman, Joseph; Pearlstein, Arne J.; Wilcutts, Mark A.

    1990-01-01

    This paper presents a control-oriented model of the flow, reaction, and transport processes in liquid propellant rocket combustion chambers, based on the multicomponent conservation laws of gas dynamics. This model provides a framework for the inclusion of detailed chemical kinetic relations, viscous and other dissipative effects, a variety of actuators and sensors, as well as process and measurement disturbances. In addition to its potential usefulness to the designer in understanding the dynamical complexity of the system and the sources of model uncertainty, the model provides a rigorous basis for control system design. An appraisal of current and feasible actuators and sensors, and their mathematical representation are included.

  12. Nonlinear analysis of longitudinal-mode liquid propellant rocket combustion instability

    Science.gov (United States)

    Jeng, San-Mou; Litchford, Ron J.

    1991-01-01

    A computational technique for the nonlinear analysis of longitudinal-mode liquid propellant rocket combustion instability is examined based on the unsteady, quasi-one-dimensional Euler equations with appropriate source terms introduced to account for interphase transport coupling with the spray. The method is first assessed for unsteady, nonreacting, isentropic duct flow with specified admittances at the outflow and inflow boundaries. For small amplitude disturbances, numerical results based on a sufficient number of grid points compare favorably with predictions from a small-disturbance linearized analysis.

  13. Prediction of high frequency combustion instability in liquid propellant rocket engines

    Science.gov (United States)

    Kim, Y. M.; Chen, C. P.; Ziebarth, J. P.; Chen, Y. S.

    1992-01-01

    The present use of a numerical model developed for the prediction of high-frequency combustion stabilities in liquid propellant rocket engines focuses on (1) the overall behavior of nonlinear combustion instabilities (2) the effects of acoustic oscillations on the fuel-droplet vaporization and combustion process in stable and unstable engine operating conditions, oscillating flowfields, and liquid-fuel trajectories during combustion instability, and (3) the effects of such design parameters as inlet boundary conditions, initial spray conditions, and baffle length. The numerical model has yielded predictions of the tangential-mode combustion instability; baffle length and droplet size variations are noted to have significant effects on engine stability.

  14. Contact diagnostics of combustion products of rocket engines, their units, and systems

    Science.gov (United States)

    Ivanov, N. N.; Ivanov, A. N.

    2013-12-01

    This article is devoted to a new block-module device used in the diagnostics of condensed combustion products of rocket engines during research and development with liquid-propellant rocket engines (Glushko NPO Energomash; engines RD-171, RD-180, and RD-191) and solid-propellant rocket motors. Soot samplings from the supersonic high-temperature jet of a high-power liquid-propellant rocket engine were taken by the given device for the first time in practice for closed-exhaust lines. A large quantity of significant results was also obtained during a combustion investigation of solid propellants within solid-propellant rocket motors.

  15. Unsteady Processes in Solid Propellant Combustion,

    Science.gov (United States)

    1977-05-01

    of a burning solid to ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~ 10. — very high frecuency pressure oscillations 1 . When the

  16. Green Propellants

    OpenAIRE

    Rahm, Martin

    2010-01-01

    To enable future environmentally friendly access to space by means of solid rocket propulsion a viable replacement to the hazardous ammonium perchlorate oxidizer is needed. Ammonium dinitramide (ADN) is one of few such compounds currently known. Unfortunately compatibility issues with many polymer binder systems and unexplained solid-state behavior have thus far hampered the development of ADN-based propellants. Chapters one, two and three offer a general introduction to the thesis, and into ...

  17. Numerical Exploration of Solid Rocket Motor Blast Tube Flow Field

    OpenAIRE

    Afroz Javed; Sinha, P.K.; Debasis Chakraborty

    2013-01-01

    The blast tube flowfield of a solid rocket motor is explored numerically by solving 3-D RANS equations with SST Turbulence model using a commercial computational fluid dynamics (CFD) software CFX-10. Parametric studies are carried out to find out the effect of the blast tube diameter on the total pressure loss in the rocket motor. It is observed that the total pressure loss in the rocket motor is less than 4 per cent and the blast tube is contributing less than 1 per cent. It is also found ou...

  18. Scale Effects on Quasi-Steady Solid Rocket Internal Ballistic Behaviour

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2010-11-01

    Full Text Available The ability to predict with some accuracy a given solid rocket motor’s performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size. The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor’s internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor’s internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale.

  19. Scale Effects on Quasi-Steady Solid Rocket Internal Ballistic Behaviour

    Energy Technology Data Exchange (ETDEWEB)

    Greatrix, David R. [Department of Aerospace Engineering, Ryerson University, 350 Victoria Street, Toronto, Ontario (Canada)

    2010-10-15

    The ability to predict with some accuracy a given solid rocket motor’s performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size). The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor’s internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor’s internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale.

  20. Scale effects on quasi-steady solid rocket internal ballistic behaviour

    Energy Technology Data Exchange (ETDEWEB)

    Greatrix, D. R. [Department of Aerospace Engineering, Ryerson University, 350 Victoria Street, Toronto, Ontario, M5B2K3 (Canada)

    2010-11-15

    The ability to predict with some accuracy a given solid rocket motor's performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size). The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor's internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor's internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale. (author)

  1. Investigation of flame driving and flow turning in axial solid rocket instabilities

    Science.gov (United States)

    Zinn, Ben T.; Daniel, Brady R.; Matta, Lawrence M.

    1993-08-01

    An understanding of the processes responsible for driving and damping acoustic oscillations in solid rocket motors is necessary for developing practical design methods that eliminate or reduce the occurrence combustion instabilities. While state of the art solid rocket stability prediction methods generally account for the flow turning loss, the magnitude and characteristics of this loss have never been fully investigated. Results of an investigation of the role of the flow turning loss in the stability of solid rockets and its dependence upon motor design and operating parameters are described. A one dimensional acoustic stability equation that verifies that the flow turning loss term is appropriately included in the one dimensional stability formulation was derived for a chamber with a constant mean temperature and pressure by an approach independent from that of Culick. This study was extended providing the background and expressions needed to guide an experimental study of the flow turning loss in the presence of mean temperature and density gradients. This allows the study of combustion systems in which mean temperature gradients and heat losses are significant. The relevant conservation equations were solved numerically for the experimental configuration in order to predict the behavior of the flow turning loss and to assist in the analysis of experimental results. Experiments performed, with and without combustion, showed that the flow turning loss strongly depends upon the propellant burning rate and the location of the flow turning region relative to the standing pressure wave.

  2. Ignition of a polymer propellant of hybrid rocket motor by a hot particle

    Science.gov (United States)

    Glushkov, D. O.; Kuznetsov, G. V.; Strizhak, P. A.

    2017-04-01

    The ignition of polymethylmethacrylate (typical model propellant of the hybrid rocket motor) by a hot particle in a shape of parallelepiped, polyhedron, disk is investigated numerically. The initial temperature of a heat source varied within the range 950-1150 K, size of particle - within the range 2-6 mm. It is established that varying these parameters influenced significantly the main characteristic of the process - ignition delay time under ignition conditions close to critical. For considered shape of particles, ignition delay time is in ascending sequence: parallelepiped, polyhedron, disk. Three polymer ignition regimes, which characterized by the initial temperature of a heat source, ignition delay time and a location of an ignition zone in a vicinity of a hot particle, are emphasized. It is illustrated that taking into account the dependence of thermal and physical characteristics of polymethylmethacrylate on temperature, the ignition delay time increased due to augmentation of energy accumulated by a subsurface layer.

  3. Thermo-mechanical concepts applied to modeling liquid propellant rocket engine stability

    Science.gov (United States)

    Kassoy, David R.; Norris, Adam

    2016-11-01

    The response of a gas to transient, spatially distributed energy addition can be quantified mathematically using thermo-mechanical concepts available in the literature. The modeling demonstrates that the ratio of the energy addition time scale to the acoustic time scale of the affected volume, and the quantity of energy added to that volume during the former determine the whether the responses to heating can be described as occurring at nearly constant volume, fully compressible or nearly constant pressure. Each of these categories is characterized by significantly different mechanical responses. Application to idealized configurations of liquid propellant rocket engines provides an opportunity to identify physical conditions compatible with gasdynamic disturbances that are sources of engine instability. Air Force Office of Scientific Research.

  4. Inverse synthetic aperture radar imagery of a man with a rocket propelled grenade launcher

    Science.gov (United States)

    Tran, Chi N.; Innocenti, Roberto; Kirose, Getachew; Ranney, Kenneth I.; Smith, Gregory

    2004-08-01

    As the Army moves toward more lightly armored Future Combat System (FCS) vehicles, enemy personnel will present an increasing threat to U.S. soldiers. In particular, they face a very real threat from adversaries using shoulder-launched, rocket propelled grenade (RPG). The Army Research Laboratory has utilized its Aberdeen Proving Ground (APG) turntable facility to collect very high resolution, fully polarimetric Ka band radar data at low depression angles of a man holding an RPG. In this paper, we examine the resulting low resolution and high resolution range profiles; and based on the observed radar cross section (RCS) value, we attempt to determine the utility of Ka band radar for detecting enemy personnel carrying RPG launchers.

  5. Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

    Science.gov (United States)

    Yang, H. Q.; West, Jeff; Harris, Robert E.

    2014-01-01

    Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor

  6. Modeling of Nonlinear Combustion Instability in Solid Propellant Rocket Motors

    Science.gov (United States)

    1984-02-01

    techniques required to deal with such discontinuities are discussed in Sections 2.5 and 6.4. Ihe inclusion of end burning effects does not require any...provide a vent time of approximately 20% of the period of the fundamento acoustic mode of the chamber. By varying the design of the pulser, it is possible

  7. Behavior Of Joint Seal In Solid Rocket Booster

    Science.gov (United States)

    Moore, Carleton J.

    1988-01-01

    Report analyzes behavior of O-ring seals in case of Solid Rocket Booster. Numerical simulations of transient response of seal presented with measurements of relevant mechanical properties of O-rings to show there is range of operating conditions in which seal can fail.

  8. Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion Measurements

    Science.gov (United States)

    Palaszewski, Bryan; Zakany, James S.

    1996-01-01

    A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face

  9. Elastomeric Thermal Insulation Design Considerations in Long, Aluminized Solid Rocket Motors

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

    An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.

  10. Numerical Simulation of Multiphase Flow in Solid Rocket Motors

    Science.gov (United States)

    Attili, A.; Favini, B.; Di Giacinto, M.; Serraglia, F.

    2009-01-01

    In the paper a general mathematical description of the flow in the internal chamber of solid rocket motors is presented. The formulation adopted take into account the multi-species and multiphase, reactive, multidimensional characteristics of the flow. The grain combustion is described by a pressure dependent law; aluminum droplet are modelled by a Lagrangian approach, coupled with the Eulerian formulation adopted for the gas phase. The mathematical model has been implemented in a simulation code and several simulations have been performed; in particular in the paper the re- sults for two geometries are described: a simple cylindrical port-area rocket and the Zefiro 9 SRM.

  11. Acoustic-mean flow interaction in solid rocket motors using Navier-Stokes equations

    Science.gov (United States)

    Vuillot, F.; Avalon, G.

    1988-07-01

    The presented numerical solution of laminar, two-dimensional, compressible and unsteady Navier-Stokes equations is aimed at a complete description of acoustic boundary layers that develop above a burning propellant. Such acoustic boundary layers are responsible for the so-called flow turning losses and also govern the local unsteady flow conditions that are seen by the burning propellant and to which it finally responds. In those respects, a complete understanding of such acoustic boundary layers is essential to improve existing solid rocket stability prediction codes. The full numerical solution of the Navier-Stokes equations permits to naturally incorporate into the analysis all the features of two-dimensional rocket chamber mean flow field. After a standing wave pattern is established through forcing at a given frequency, a special Fourier treatment is used to put the numerical results in a form directly comparable to available linear acoustic data. The presented results indicate that the acoustic boundary layer is substantially thinner than predicted by simplified models. Moreover, its acoustic admittance is found to significantly vary along the chamber, a result that is of major importance to stability predictions. Finally, the acoustic field is found to be rotational over a significant volume of the chamber, leading to a volume flow turning loss, rather than to a pure surface effect as usually assumed.

  12. Development of Thermal Barriers for Solid Rocket Motor Nozzle Joints

    Science.gov (United States)

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.

    1999-01-01

    The Space Shuttle solid rocket motor case assembly joints are sealed using conventional 0-ring seals. The 5500+F combustion gases are kept a safe distance away from the seals by thick layers of insulation. Special joint-fill compounds are used to fill the joints in the insulation to prevent a direct flowpath to the seals. On a number of occasions. NASA has observed in several of the rocket nozzle assembly joints hot gas penetration through defects in the joint- fill compound. The current nozzle-to-case joint design incorporates primary, secondary and wiper (inner-most) 0-rings and polysulfide joint-fill compound. In the current design, 1 out of 7 motors experience hot gas to the wiper 0-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper 0-ring results in extensive reviews before resuming flight. NASA and solid rocket motor manufacturer Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and a thermal barrier, This paper presents burn-resistance, temperature drop, flow and resiliency test results for several types of NASA braided carbon-fiber thermal barriers. Burn tests were performed to determine the time to burn through each of the thermal barriers when exposed to the flame of an oxy-acetylene torch (5500 F), representative of the 5500 F solid rocket motor combustion temperatures. Thermal barriers braided out of carbon fibers endured the flame for over 6 minutes, three times longer than solid rocket motor burn time. Tests were performed on two thermal barrier braid architectures, denoted Carbon-3 and Carbon-6, to measure the temperature drop across and along the barrier in a compressed state when subjected to the flame of an oxyacetylene torch. Carbon-3 and Carbon-6 thermal barriers were excellent insulators causing temperature drops through their diameter of up to a 2800 and 2560 F. respectively. Gas temperature 1/4" downstream of the thermal barrier were within the

  13. Bonding agensi za kompozitna raketna goriva / Bonding agents for composite rocket propellants

    Directory of Open Access Journals (Sweden)

    Mirjana Petrić

    2003-01-01

    Full Text Available U radu je analiziran uticaj bonding agensd na mehaničke karakteristike i viskozitet kompozitnih raketnih goriva na bazi hidroksiterminiranog polibutadiena i toluendiizocijanata, odnosno izoforondiizocijanata. Komparirana je efikasnost bonding agensd (BA različite strukture trietilentetramina (TET, tris-(2-metil aziridinil fosfin oksida (MAPO i N,N-di (2-hidroksietil-4,4-dimetilhidantoina (DHE. Efikasnost delovanja bonding agensa utvrdenaje na osnovu parametara dobijenih iz testa jednoosnog zatezanja uzoraka (σm - zatezna čvrstoća pri maksimalnoj sili, εm relativno izduženje pri maksimalnoj sili i εp - relativno izduženje pri prekidu. Mehaničke karakteristike merene su u temperaturskom području od -50°C do 50°C. Analiziran je uticaj BA na brzinu promene viskoziteta goriva. / The paper analyzes the effects of bonding agents on mechanical properties and viscosity of composite rocket propellants based on hydroxiterminated polybutadiene and toluendiizocyanate, i.e. izophorondiizocyanate. The efficiency of bonding agents (BA of different structure has been compared including triethylentetramine (TET, tris-(2-methyl azyrinidile phosphine oxide (MAPO and N, N-di (2-hydroxiethyl-4,4-dimethylhydantoine (DHE. The BA efficiency has been determined on the basis of parameters obtained by uniaxial tensile tests (σm - tensile strength at maximum force, εm - relative allongation at maximum force and εp - relative allongation at fracture. The mechanical properties have been measured at the temperature range from -50°C to 50°C. The BA effects on propellant viscosity change rates have been analyzed.

  14. The identification of parameter of solid rocket motor

    Science.gov (United States)

    Fang, Dingyou

    1992-08-01

    An end-burning grain in which long metal wires are embedded is used here as an example of parameter identification in solid rocket motors. The results indicate that when the deposition and erosion at the nozzle throat are not considered or the characteristic velocity is inaccurate or the identification technique of identification and parameter initial value identification are inaccurate, the identification results are all inaccurate. Parameters measured and information obtained experimentally can be used to check the credibility of identified parameters.

  15. Space shuttle program solid rocket booster decelerator subsystem

    Science.gov (United States)

    Barnard, J. W.

    1985-01-01

    The recovery of the Solid Rocket Boosters presented a major challenge. The SRB represents the largest payload ever recovered and presents the added complication that it is continually emitting hot gases and burning particles of insulation and other debris. Some items, such as portions of the nozzle, are large enough to burn through the nylon parachute material. The SRB Decelerator Subsystem program was highly successful in that no SRB has been lost as a result of inadequate performance of the DSS.

  16. Hybrid Rocket Technology

    OpenAIRE

    Sankaran Venugopal; K K Rajesh; V. Ramanujachari

    2011-01-01

    With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems. Classical hybrids can be throttled for thrust tailoring, perform in-flight motor shutdown and restart. In classical hybrids, the fuel is stored in the form of a solid grain, requiring only half the feed system hardware of liquid bipropellant engines. The commonly used fuels are benign, nonto...

  17. The measurement, modeling, and prediction of traction for rocket propellant 1

    Science.gov (United States)

    Tevaarwerk, J. L.

    1989-01-01

    Traction tests were performed on RP-1, a common kerosene based rocket propellant. Traction data on this fluid are required for purposes of turbopump bearing design, using codes such as SHABERTH. To obtain the traction data, an existing twin disc machine was used, operating under the side slip mode and using elliptical contacts. The range of test variables were: contact peak Hertz stress from 1.0 to 2.0 GPa, disc surface speed from 10 to 50 m/s, fluid inlet temperature from 30 to 70 C, and with a contact aspect ratio of 1.7. The resulting traction curves were reduced to fundamental fluid property parameters using the Johnson and Tevaarwerk traction model. Theoretical traction predictions were performed by back substitution of the fundamental properties into the traction model. Comparison of the predicted with the measured curves gives a high degree of confidence in the correctness of the traction model. For purposes of input to the NASA SHABERTH program, the traction model was next used to predict the expected traction of RP-1 under line contact conditions.

  18. Modular simulation software development for liquid propellant rocket engines based on MATLAB Simulink

    Directory of Open Access Journals (Sweden)

    Naderi Mahyar

    2017-01-01

    Full Text Available Focusing on Liquid Propellant Rocket Engine (LPRE major components, a steady state modular simulation software has been established in MATLAB Simulink. For integrated system analysis, a new algorithm dependant on engine inlet mass flow rate and pressure is considered. Using the suggested algorithm, it is possible to evaluate engine component operation, similar to the known initial parameters during hot fire test of the engine on stand. As a case study, the reusable Space Shuttle Main Engine (SSME has been selected and the simulation has been performed to predict engine’s throttled operation at 109 percent of the nominal thrust value. For this purpose the engine actual flow diagram has been converted to 34 numerical modules and the engine has been modelled by solving a total of 101 steady state mathematic equations. The mean error of the simulation results is found to be less than 5% compared with the published SSME data. Using the presented idea and developed modules, it is possible to build up the numerical model and simulate other LPREs.

  19. Baseline Computational Fluid Dynamics Methodology for Longitudinal-Mode Liquid-Propellant Rocket Combustion Instability

    Science.gov (United States)

    Litchford, R. J.

    2005-01-01

    A computational method for the analysis of longitudinal-mode liquid rocket combustion instability has been developed based on the unsteady, quasi-one-dimensional Euler equations where the combustion process source terms were introduced through the incorporation of a two-zone, linearized representation: (1) A two-parameter collapsed combustion zone at the injector face, and (2) a two-parameter distributed combustion zone based on a Lagrangian treatment of the propellant spray. The unsteady Euler equations in inhomogeneous form retain full hyperbolicity and are integrated implicitly in time using second-order, high-resolution, characteristic-based, flux-differencing spatial discretization with Roe-averaging of the Jacobian matrix. This method was initially validated against an analytical solution for nonreacting, isentropic duct acoustics with specified admittances at the inflow and outflow boundaries. For small amplitude perturbations, numerical predictions for the amplification coefficient and oscillation period were found to compare favorably with predictions from linearized small-disturbance theory as long as the grid exceeded a critical density (100 nodes/wavelength). The numerical methodology was then exercised on a generic combustor configuration using both collapsed and distributed combustion zone models with a short nozzle admittance approximation for the outflow boundary. In these cases, the response parameters were varied to determine stability limits defining resonant coupling onset.

  20. Introduction of laser initiation for the 48-inch Advanced Solid Rocket Motor (ASRM) test motors at Marshall Space Flight Center (MSFC)

    Science.gov (United States)

    Zimmerman, Chris J.; Litzinger, Gerald E.

    1993-01-01

    The Advanced Solid Rocket Motor is a new design for the Space Shuttle Solid Rocket Booster. The new design will provide more thrust and more payload capability, as well as incorporating many design improvements in all facets of the design and manufacturing process. A 48-inch (diameter) test motor program is part of the ASRM development program. This program has multiple purposes for testing of propellent, insulation, nozzle characteristics, etc. An overview of the evolution of the 48-inch ASRM test motor ignition system which culminated with the implementation of a laser ignition system is presented. The laser system requirements, development, and operation configuration are reviewed in detail.

  1. Multisized Inert Particle Loading for Solid Rocket Axial Combustion Instability Suppression

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2012-01-01

    Full Text Available In the present investigation, various factors and trends, related to the usage of two or more sets of inert particles comprised of the same material (nominally aluminum but at different diameters for the suppression of axial shock wave development, are numerically predicted for a composite-propellant cylindrical-grain solid rocket motor. The limit pressure wave magnitudes at a later reference time in a given pulsed firing simulation run are collected for a series of runs at different particle sizes and loading distributions and mapped onto corresponding attenuation trend charts. The inert particles’ presence in the central core flow is demonstrated to be an effective means of instability symptom suppression, in correlating with past experimental successes in the usage of particles. However, the predicted results of this study suggest that one needs to be careful when selecting more than one size of particle for a given motor application.

  2. Reusable Solid Rocket Motor - Accomplishments, Lessons, and a Culture of Success

    Science.gov (United States)

    Moore, Dennis R.; Phelps, Willie J.

    2011-01-01

    The Reusable Solid Rocket Motor represents the largest solid rocket motor ever flown and the only human rated solid motor. Each Reusable Solid Rocket Motor (RSRM) provides approximately 3-million lb of thrust to lift the integrated Space Shuttle vehicle from the launch pad. The motors burn out approximately 2 minutes later, separate from the vehicle and are recovered and refurbished. The size of the motor and the need for high reliability were challenges. Thrust shaping, via shaping of the propellant grain, was needed to limit structural loads during ascent. The motor design evolved through several block upgrades to increase performance and to increase safety and reliability. A major redesign occurred after STS-51L with the Redesigned Solid Rocket Motor. Significant improvements in the joint sealing systems were added. Design improvements continued throughout the Program via block changes with a number of innovations including development of low temperature o-ring materials and incorporation of a unique carbon fiber rope thermal barrier material. Recovery of the motors and post flight inspection improved understanding of hardware performance, and led to key design improvements. Because of the multidecade program duration material obsolescence was addressed, and requalification of materials and vendors was sometimes needed. Thermal protection systems and ablatives were used to protect the motor cases and nozzle structures. Significant understanding of design and manufacturing features of the ablatives was developed during the program resulting in optimization of design features and processing parameters. The project advanced technology in eliminating ozone-depleting materials in manufacturing processes and the development of an asbestos-free case insulation. Manufacturing processes for the large motor components were unique and safety in the manufacturing environment was a special concern. Transportation and handling approaches were also needed for the large

  3. Internal Flow Simulation of Enhanced Performance Solid Rocket Booster for the Space Transportation System

    Science.gov (United States)

    Ahmad, Rashid A.; McCool, Alex (Technical Monitor)

    2001-01-01

    An enhanced performance solid rocket booster concept for the space shuttle system has been proposed. The concept booster will have strong commonality with the existing, proven, reliable four-segment Space Shuttle Reusable Solid Rocket Motors (RSRM) with individual component design (nozzle, insulator, etc.) optimized for a five-segment configuration. Increased performance is desirable to further enhance safety/reliability and/or increase payload capability. Performance increase will be achieved by adding a fifth propellant segment to the current four-segment booster and opening the throat to accommodate the increased mass flow while maintaining current pressure levels. One development concept under consideration is the static test of a "standard" RSRM with a fifth propellant segment inserted and appropriate minimum motor modifications. Feasibility studies are being conducted to assess the potential for any significant departure in component performance/loading from the well-characterized RSRM. An area of concern is the aft motor (submerged nozzle inlet, aft dome, etc.) where the altered internal flow resulting from the performance enhancing features (25% increase in mass flow rate, higher Mach numbers, modified subsonic nozzle contour) may result in increased component erosion and char. To assess this issue and to define the minimum design changes required to successfully static test a fifth segment RSRM engineering test motor, internal flow studies have been initiated. Internal aero-thermal environments were quantified in terms of conventional convective heating and discrete phase alumina particle impact/concentration and accretion calculations via Computational Fluid Dynamics (CFD) simulation. Two sets of comparative CFD simulations of the RSRM and the five-segment (IBM) concept motor were conducted with CFD commercial code FLUENT. The first simulation involved a two-dimensional axi-symmetric model of the full motor, initial grain RSRM. The second set of analyses

  4. Health Monitoring and Diagnosis of Solid Rocket Motors with Bore Cracks

    Science.gov (United States)

    2015-11-01

    Technical Paper 3. DATES COVERED (From - To) January 2014-February 2015 4. TITLE AND SUBTITLE Health Monitoring and Diagnosis of Solid Rocket Motors with... rocket motors at various storage temperatures. Capabilities of a rocket motor health monitoring system are assessed based on the assumption that the...system can detect critical bore cracks in solid rocket motors. 15. SUBJECT TERMS structural health monitoring (SHM) · structural integrity · damage

  5. Performance Investigation of Solid-Rocket Motor with Nozzle Throat Erosion

    OpenAIRE

    Suwicha Chankapoe; Nattawat Winya; Narupon Pittayaprasertkul

    2013-01-01

    In order to determine the performance and key design parameters of rocket, the erosion of nozzle throat during solid rocket motor burning have to be calculated. This study aims to predict the nozzle throat erosion in solid rocket motors according to the thrust profile of motor in operating conditions and develop a model for optimum performance of rocket. We investigate the throat radius change in the static test programs. The standard method and thrust coefficient are used for adjusting into...

  6. Numerical Evaluation of the Use of Aluminum Particles for Enhancing Solid Rocket Motor Combustion Stability

    Directory of Open Access Journals (Sweden)

    David Greatrix

    2015-02-01

    Full Text Available The ability to predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave development in the motor chamber flow. With the focus of the present study placed on reactive particles, a numerical internal ballistic model incorporating relevant elements, such as a transient, frequency-dependent combustion response to axial pressure wave activity above the burning propellant surface, is applied to the investigation of using aluminum particles within the central internal flow (particles whose surfaces nominally regress with time, as a function of current particle size, as they move downstream as a means of suppressing instability-related symptoms in a cylindrical-grain motor. The results of this investigation reveal that the loading percentage and starting size of the aluminum particles have a significant influence on reducing the resulting transient pressure wave magnitude.

  7. Study of plasticizer diffusion in a solid rocket motor´s bondline

    Directory of Open Access Journals (Sweden)

    Juliano Libardi

    2009-06-01

    Full Text Available This work aims to determine the diffusion coefficient of the plasticizers dibutyl phthalate (DBP, dioctyl phthalate (DOP and dioctyl azelate (DOZ on the internal insulating layer of solid rocket motors. These plasticizers are originally present in the layers of rubber, liner and propellant, respectively. This species are not chemically bonded and tend to diffuse from propellant to insulating and vice versa. A computer program based on the mathematical model of Fick’s second Law of diffusion was developed to perform the calculus from the concentration data obtained by gas chromatographic (GC analyses. The samples were prepared with two different adhesive liners; one conventional (LHNA and the other with barrier properties (LHNT. A common feature of both liners was that they were synthesized by the reaction of hydroxyl-terminated polybutadiene (HTPB and diisocyanates. However, a bond promoter was used to increase the crosslink density of the LHNT liner and to improve its performance as barrier against the diffusion. The effects of the diffusion of the plasticizers were also investigated by hardness analyses, which were executed on samples aged at room temperature and at 80ºC. The results showed an increase trend for the samples aged at room temperature and an opposite behavior for the tests carried out at 80ºC.

  8. Fluid-Structure Interaction Effects on Mass Flow Rates in Solid Rocket Motors

    Science.gov (United States)

    2015-09-02

    in Solid Rocket Motors 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) William Harrigan 5d. PROJECT NUMBER...Determination of mass flow rate in a solid rocket motor is critical in the design of a new motor due to its effect on the thrust produced. Fluid...mass flow rates. The FSI analyses with two‐way coupling provided a more accurate assessment of solid rocket motor internal ballistics. 15. SUBJECT

  9. Inter-Batch Variation and the Effect of Casting Vacuum on Ballistic and Mechanical Properties of a High Performing Cast Composite Rocket Propellant

    Science.gov (United States)

    2014-12-01

    Tensile testing of propellant in the form of ‘ dog bones’ or uniaxial tensile test pieces cut from cast slabs of propellant was performed to assess inter... dog bones were cut from the top of the mould. It is possible that the strands for both partial and full vacuum had a slightly higher solids content... Behaviour of AP/HTPB Propellant Using Coarse and Fine AP Particles. Propellants, Explosives, Pyrotechnics, 2011. 36: p. 57-64. 9. Smith, P.C

  10. Direct Electrical Arc Ignition of Hybrid Rocket Motors

    OpenAIRE

    Judson, Michael I., Jr.

    2015-01-01

    Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven dicult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintrod...

  11. Concerning the problem of dynamic damping of the vibration combustion self-oscillations in a liquid-propellant rocket engine

    Science.gov (United States)

    Basok, B. I.; Gotsulenko, V. V.; Gotsulenko, V. N.

    2012-11-01

    The reason for the decrease in the amplitude of longitudinal vibration combustion self-oscillations in the combustion chamber of a liquid-propellant rocket engine by means of antipulse partitions has been justified. A mathematical model of the development of combustion instability in such a chamber on attachment of a Helmholtz resonator to it has been obtained. The character of the damping of vibration combustion self-oscillations excited by the action of the Crocco mechanisms and negative thermal resistance, when varying the acoustic parameters of the resonator and of the pressure head characteristics of combustion chamber is established.

  12. Combustion of single and agglomerated aluminum particles in solid rocket motor flows

    Science.gov (United States)

    Melcher, John Charles, IV

    2001-07-01

    Single and agglomerated aluminum droplets were studied in a solid rocket motor (SRM) test chamber with optical access to the internal flow at 6--22 atm and 2300 K. The chamber was pressurized by burning a main grain AP/HTPB propellant, and the burning aluminum droplets were generated by a smaller aluminized solid propellant sample, center-mounted in the flow. A 35 mm camera was used with a chopper wheel to give droplet flame diameter vs. time measurements of the burning droplets in flight, from which bum-rate laws were developed. A high-speed video CCD was used with high-magnification optics in order to image the flame/smoke cloud surrounding the burning liquid droplets. The intensity profiles of the droplet images were de-convoluted using an Abel inversion to give true intensity profiles. Both single and agglomerated droplets were studied, where agglomerates are comprised of hundreds of parent particles or more. The Abel inversion results show that the relative smoke cloud size is not constant with diameter, but instead grows as the droplet shrinks, by ˜D -0.5, for both the single and agglomerated droplets. Measured diameter trajectories show that for single droplets, the diameter law is D 0.75 = DO0.75 = 8·t [mu m, msec], and for agglomerated droplets, D 1.0 = Do1.0 - 20·t, such that the single droplets burn faster than the agglomerates. For both single and agglomerated droplets, the burning rate slope k did not change significantly over the chamber pressure studied. Lastly, a model was developed to describe the oxide cap accumulation on the droplet surface from the oxide smoke cloud surrounding the droplet. Results suggest that less oxide accumulates in high-pressure SRMs when considering mass burning rates for different relative cap sizes. The thermophoretic force, which can control oxide transport only over the cap, decreases with pressure.

  13. Ecological effects and environmental fate of solid rocket exhaust

    Science.gov (United States)

    Nimmo, B.; Stout, I. J.; Mickus, J.; Vickers, D.; Madsen, B.

    1974-01-01

    Specific target processes were classified as to the chemical, chemical-physical, and biological reactions and toxic effects of solid rocket emissions within selected ecosystems at Kennedy Space Center. Exposure of Citris seedlings, English peas, and bush beans to SRM exhaust under laboratory conditions demonstrated reduced growth rates, but at very high concentrations. Field studies of natural plant populations in three diverse ecosystems failed to reveal any structural damage at the concentration levels tested. Background information on elemental composition of selected woody plants from two terrestrial ecosystems is reported. LD sub 50 for a native mouse (peromysous gossypinus) exposed to SRM exhaust was determined to be 50 ppm/g body weight. Results strongly indicate that other components of the SRM exhaust act synergically to enhance the toxic effects of HCl gas when inhaled. A brief summary is given regarding the work on SRM exhaust and its possible impact on hatchability of incubating bird eggs.

  14. Indirect and direct methods for measuring a dynamic throat diameter in a solid rocket motor

    Science.gov (United States)

    Colbaugh, Lauren

    In a solid rocket motor, nozzle throat erosion is dictated by propellant composition, throat material properties, and operating conditions. Throat erosion has a significant effect on motor performance, so it must be accurately characterized to produce a good motor design. In order to correlate throat erosion rate to other parameters, it is first necessary to know what the throat diameter is throughout a motor burn. Thus, an indirect method and a direct method for determining throat diameter in a solid rocket motor are investigated in this thesis. The indirect method looks at the use of pressure and thrust data to solve for throat diameter as a function of time. The indirect method's proof of concept was shown by the good agreement between the ballistics model and the test data from a static motor firing. The ballistics model was within 10% of all measured and calculated performance parameters (e.g. average pressure, specific impulse, maximum thrust, etc.) for tests with throat erosion and within 6% of all measured and calculated performance parameters for tests without throat erosion. The direct method involves the use of x-rays to directly observe a simulated nozzle throat erode in a dynamic environment; this is achieved with a dynamic calibration standard. An image processing algorithm is developed for extracting the diameter dimensions from the x-ray intensity digital images. Static and dynamic tests were conducted. The measured diameter was compared to the known diameter in the calibration standard. All dynamic test results were within +6% / -7% of the actual diameter. Part of the edge detection method consists of dividing the entire x-ray image by an average pixel value, calculated from a set of pixels in the x-ray image. It was found that the accuracy of the edge detection method depends upon the selection of the average pixel value area and subsequently the average pixel value. An average pixel value sensitivity analysis is presented. Both the indirect

  15. Advanced Multi-Phase Flow CFD Model Development for Solid Rocket Motor Flowfield Analysis

    Science.gov (United States)

    Liaw, Paul; Chen, Y. S.; Shang, H. M.; Doran, Denise

    1993-01-01

    It is known that the simulations of solid rocket motor internal flow field with AL-based propellants require complex multi-phase turbulent flow model. The objective of this study is to develop an advanced particulate multi-phase flow model which includes the effects of particle dynamics, chemical reaction and hot gas flow turbulence. The inclusion of particle agglomeration, particle/gas reaction and mass transfer, particle collision, coalescence and breakup mechanisms in modeling the particle dynamics will allow the proposed model to realistically simulate the flowfield inside a solid rocket motor. The Finite Difference Navier-Stokes numerical code FDNS is used to simulate the steady-state multi-phase particulate flow field for a 3-zone 2-D axisymmetric ASRM model and a 6-zone 3-D ASRM model at launch conditions. The 2-D model includes aft-end cavity and submerged nozzle. The 3-D model represents the whole ASRM geometry, including additional grain port area in the gas cavity and two inhibitors. FDNS is a pressure based finite difference Navier-Stokes flow solver with time-accurate adaptive second-order upwind schemes, standard and extended k-epsilon models with compressibility corrections, multi zone body-fitted formulations, and turbulence particle interaction model. Eulerian/Lagrangian multi-phase solution method is applied for multi-zone mesh. To simulate the chemical reaction, penalty function corrected efficient finite-rate chemistry integration method is used in FDNS. For the AL particle combustion rate, the Hermsen correlation is employed. To simulate the turbulent dispersion of particles, the Gaussian probability distribution with standard deviation equal to (2k/3)(exp 1/2) is used for the random turbulent velocity components. The computational results reveal that the flow field near the juncture of aft-end cavity and the submerged nozzle is very complex. The effects of the turbulent particles affect the flow field significantly and provide better

  16. A stochastic pocket model for aluminum agglomeration in solid propellants

    Energy Technology Data Exchange (ETDEWEB)

    Gallier, Stany [SNPE Materiaux Energetiques, Vert le Petit (France)

    2009-04-15

    A new model is derived to estimate the size and fraction of aluminum agglomerates at the surface of a burning propellant. The basic idea relies on well-known pocket models in which aluminum is supposed to aggregate and melt within pocket volumes imposed by largest oxidizer particles. The proposed model essentially relaxes simple assumptions of previous pocket models on propellant structure by accounting for an actual microstructure obtained by packing. The use of statistical tools from stochastic geometry enables to determine a statistical pocket size volume and hence agglomerate diameter and agglomeration fraction. Application to several AP/Al propellants gives encouraging results that are shown to be superior to former pocket models. (Abstract Copyright [2009], Wiley Periodicals, Inc.)

  17. Numerical and experimental study of liquid breakup process in solid rocket motor nozzle

    Science.gov (United States)

    Yen, Yi-Hsin

    Rocket propulsion is an important travel method for space exploration and national defense, rockets needs to be able to withstand wide range of operation environment and also stable and precise enough to carry sophisticated payload into orbit, those engineering requirement makes rocket becomes one of the state of the art industry. The rocket family have been classified into two major group of liquid and solid rocket based on the fuel phase of liquid or solid state. The solid rocket has the advantages of simple working mechanism, less maintenance and preparing procedure and higher storage safety, those characters of solid rocket make it becomes popular in aerospace industry. Aluminum based propellant is widely used in solid rocket motor (SRM) industry due to its avalibility, combusion performance and economical fuel option, however after aluminum react with oxidant of amonimum perchrate (AP), it will generate liquid phase alumina (Al2O3) as product in high temperature (2,700˜3,000 K) combustion chamber enviornment. The liquid phase alumina particles aggromorate inside combustion chamber into larger particle which becomes major erosion calprit on inner nozzle wall while alumina aggromorates impinge on the nozzle wall surface. The erosion mechanism result nozzle throat material removal, increase the performance optimized throat diameter and reduce nozzle exit to throat area ratio which leads to the reduction of exhaust gas velocity, Mach number and lower the propulsion thrust force. The approach to avoid particle erosion phenomenon taking place in SRM's nozzle is to reduce the alumina particle size inside combustion chamber which could be done by further breakup of the alumina droplet size in SRM's combustion chamber. The study of liquid breakup mechanism is an important means to smaller combustion chamber alumina droplet size and mitigate the erosion tack place on rocket nozzle region. In this study, a straight two phase air-water flow channel experiment is set up

  18. Proposal for a Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft Project

    Data.gov (United States)

    National Aeronautics and Space Administration — A new technology, the Fission Fragment Rocket Engine (FFRE), requires small amounts of readily available, energy dense, long lasting fuel, significant thrust at...

  19. DURACON - Variable Emissivity Broadband Coatings for Liquid Propellant Rocket Nozzles Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The need exists for a fast drying, robust, low gloss, black, high emissivity coating that can be applied easily on aircraft rocket nozzles and nozzle extensions....

  20. Integration of Flex Nozzle System and Electro Hydraulic Actuators to Solid Rocket Motors

    Science.gov (United States)

    Nayani, Kishore Nath; Bajaj, Dinesh Kumar

    2017-10-01

    A rocket motor assembly comprised of solid rocket motor and flex nozzle system. Integration of flex nozzle system and hydraulic actuators to the solid rocket motors are done after transportation to the required place where integration occurred. The flex nozzle system is integrated to the rocket motor in horizontal condition and the electro hydraulic actuators are assembled to the flex nozzle systems. The electro hydraulic actuators are connected to the hydraulic power pack to operate the actuators. The nozzle-motor critical interface are insulation diametrical compression, inhibition resin-28, insulation facial compression, shaft seal `O' ring compression and face seal `O' ring compression.

  1. YMCA ROCKET RAMPAGE! SUMMER CAMP

    National Research Council Canada - National Science Library

    Anonymous

    2014-01-01

    ... & Controls, sponsored the Rocket Rampagel summer camp at the YMCA in Eklton MD. On day 1, campers took Rockets 101, constructing balloon rockets and straw rockets, followed by racket manufacturing, where campers made rocket "propellant" on day 2...

  2. Solid Rocket Booster (SRB) Flight System Integration at Its Best

    Science.gov (United States)

    Wood, T. David; Kanner, Howard S.; Freeland, Donna M.; Olson, Derek T.

    2011-01-01

    The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads, environments and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. However, the in-flight data and postflight assessment process revealed the hardware was affected much more strongly than originally anticipated. Assembly and integration of the booster subsystems required acceptance testing of reused hardware components for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed

  3. An Analysis of the Orbital Distribution of Solid Rocket Motor Slag

    Science.gov (United States)

    Horstman, Matthew F.; Mulrooney, Mark

    2007-01-01

    The contribution made by orbiting solid rocket motors (SRMs) to the orbital debris environment is both potentially significant and insufficiently studied. A combination of rocket motor design and the mechanisms of the combustion process can lead to the emission of sufficiently large and numerous by-products to warrant assessment of their contribution to the orbital debris environment. These particles are formed during SRM tail-off, or the termination of burn, by the rapid expansion, dissemination, and solidification of the molten Al2O3 slag pool accumulated during the main burn phase of SRMs utilizing immersion-type nozzles. Though the usage of SRMs is low compared to the usage of liquid fueled motors, the propensity of SRMs to generate particles in the 100 m and larger size regime has caused concern regarding their contributing to the debris environment. Particle sizes as large as 1 cm have been witnessed in ground tests conducted under vacuum conditions and comparable sizes have been estimated via ground-based telescopic and in-situ observations of sub-orbital SRM tail-off events. Using sub-orbital and post recovery observations, a simplistic number-size-velocity distribution of slag from on-orbit SRM firings was postulated. In this paper we have developed more elaborate distributions and emission scenarios and modeled the resultant orbital population and its time evolution by incorporating a historical database of SRM launches, propellant masses, and likely location and time of particulate deposition. From this analysis a more comprehensive understanding has been obtained of the role of SRM ejecta in the orbital debris environment, indicating that SRM slag is a significant component of the current and future population.

  4. Realization of a Solid-Propellant based Microthruster Using Low Temperature Co-fired Ceramics

    Directory of Open Access Journals (Sweden)

    Jaya THAKUR

    2010-06-01

    Full Text Available The introduction of micro-spacecrafts in the space industry has led to the development of various micro-propulsion techniques. Microthrusters are micropropulsion devices used in a microspacecraft for precise station keeping, orbit adjustment, attitude control, drag compensation and apogee kicking. The principle of operation of a solid propellant thruster is based on the combustion of a solid energetic material stored in a microfabricated chamber. In the current work, Low-Temperature Co-fired Ceramic (LTCC technology has been used for the realization of a solid propellant based microthruster structure. Hydroxyl Terminated Poly-Butadiene/Ammonium Perchlorate (HTPB/AP is used as the propellant. It is shown that geometric and dimensional variations in design, depending on the application requirements, can be easily implemented. Preliminary testing for micro-combustion has been done to verify the basic operation of the microthruster. A thrust value of 19.5 mN has been measured.

  5. Study on the extraction of AP components in HTPB solid propellant by water/ethanol recovery

    Science.gov (United States)

    Wang, Mingqi; Du, Shiguo; Yan, Jun

    2017-10-01

    In order to improve the recovery of ammonium perchlorate (AP) components in scrapped HTPB solid propellants, the AP in HTPB solid propellants were extracted by water/ethanol mixed solvent system. The effects of extraction time, extraction temperature, volume ratio of water/ethanol, ratio of liquid to material, thickness of sample and extraction rate on AP extraction rate were discussed. The extraction results were characterized by scanning electron microscopy (SEM). The results showed that the best process parameter for extracting AP components from HTPB solid propellants: the extraction temperature was 70°C, the extraction time was 6h, the volume ratio of water/ethanol was 1:1, the ratio of liquid to material was 6:1(water/ethanol volume and sample mass ratio, ml/g) and the extraction frequency was 2 times. Under this condition, the recovery of AP was 90.5% and the purity of AP was 91.1%.

  6. Ignition models and simulation of solid propellant of thermodynamic undersea vehicle

    Science.gov (United States)

    Zhang, Jin-Jun; Qian, Zhi-Bo; Yang, Jie; Yan, Ping

    2007-06-01

    The starting characteristics of thermodynamic undersea vehicle systems are determined by the geometry, size and combustion area of solid propellants, which directly effect liquid propellant pipeline design. It is necessary to establish accurate burning models for solid propellants. Based on combustion models using powder rings and two different solid ignition grains, namely star-shaped ignition grains and stuffed ignition grains, a mathematic model of the ignition process of the propulsion system was built. With the help of Matlab, a series of calculations were made to determine the effects of different grains on ignition characteristics. The results show that stuffed ignition grain is best suited to be the ignition grain of a thermodynamic undersea vehicle system.

  7. Modeling the effects of dissolved helium pressurant on a liquid hydrogen rocket propellant tank

    Science.gov (United States)

    Richardson, I. A.; Leachman, J. W.

    2017-12-01

    A model was developed using NASA’s Generalized Fluid System Simulation Program (GFSSP) for the self-pressurization of a liquid hydrogen propellant tank due to boil-off to determine the significance of mixture non-idealities. The GFSSP model compared the tank performance for the traditional model that assumes no helium pressurant dissolves into the liquid hydrogen propellant to an updated model that accounts for dissolved helium pressurant. Traditional NASA models have been unable to account for this dissolved helium due to a lack of fundamental property information. Recent measurements of parahydrogen-helium mixtures enabled the development of the first multi-phase Equation Of State (EOS) for parahydrogen-helium mixtures. The self-pressurization GFSSP model was run assuming that the liquid propellant was pure liquid hydrogen and assuming helium dissolved into the liquid utilizing the new helium-hydrogen EOS. The analysis shows that having dissolved helium in the propellant does not have a significant effect on the tank pressurization rate for typical tank conditions (-423 °F and 30 psia).

  8. A robust ballistic design approach for the Space Shuttle Advanced Solid Rocket Motor

    Science.gov (United States)

    Eagar, M. A.; Jordan, F. W.; Stockham, L. W.

    1993-06-01

    A robust design approach has been developed for the Space Shuttle Advanced Solid Rocket Motor (ASRM) that enhances the potential for program success. This is accomplished by application of state of the art ballistic modelling techniques coupled with an aggressive contingency planning methodology. Application of this approach addresses the design challenges associated with development of the ASRM because of it's large size, high length to diameter ratio, and demanding thrust-time trace shape requirements. Advanced ballistic modelling techniques applied include deformed grain modelling, spatial burn rate mapping, erosive burning response characterization, and ballistic/structural/CFD flow-grain interactions. Model fidelity is further improved by utilizing the extensive RSRM experience base for validation proposes. In addition to this modelling approach, development of contingency plans covers the remaining prediction uncertainties, and readily allows fine tuning of the propellant grain configuration to meet design objectives after the first motor firing. This approach promises to produce an optimum flight motor design that meets all performance objectives while accommodating program development uncertainties.

  9. A Multiconstrained Ascent Guidance Method for Solid Rocket-Powered Launch Vehicles

    Directory of Open Access Journals (Sweden)

    Si-Yuan Chen

    2016-01-01

    Full Text Available This study proposes a multiconstrained ascent guidance method for a solid rocket-powered launch vehicle, which uses a hypersonic glide vehicle (HGV as payload and shuts off by fuel exhaustion. First, pseudospectral method is used to analyze the two-stage launch vehicle ascent trajectory with different rocket ignition modes. Then, constraints, such as terminal height, velocity, flight path angle, and angle of attack, are converted into the constraints within height-time profile according to the second-stage rocket flight characteristics. The closed-loop guidance method is inferred by different spline curves given the different terminal constraints. Afterwards, a thrust bias energy management strategy is proposed to waste the excess energy of the solid rocket. Finally, the proposed method is verified through nominal and dispersion simulations. The simulation results show excellent applicability and robustness of this method, which can provide a valuable reference for the ascent guidance of solid rocket-powered launch vehicles.

  10. Multiple Changes to Reusable Solid Rocket Motors, Identifying Hidden Risks

    Science.gov (United States)

    Greenhalgh, Phillip O.; McCann, Bradley Q.

    2003-01-01

    The Space Shuttle Reusable Solid Rocket Motor (RSRM) baseline is subject to various changes. Changes are necessary due to safety and quality improvements, environmental considerations, vendor changes, obsolescence issues, etc. The RSRM program has a goal to test changes on full-scale static test motors prior to flight due to the unique RSRM operating environment. Each static test motor incorporates several significant changes and numerous minor changes. Flight motors often implement multiple changes simultaneously. While each change is individually verified and assessed, the potential for changes to interact constitutes additional hidden risk. Mitigating this risk depends upon identification of potential interactions. Therefore, the ATK Thiokol Propulsion System Safety organization initiated the use of a risk interaction matrix to identify potential interactions that compound risk. Identifying risk interactions supports flight and test motor decisions. Uncovering hidden risks of a full-scale static test motor gives a broader perspective of the changes being tested. This broader perspective compels the program to focus on solutions for implementing RSRM changes with minimal/mitigated risk. This paper discusses use of a change risk interaction matrix to identify test challenges and uncover hidden risks to the RSRM program.

  11. Reusable Solid Rocket Motor Nozzle Joint-4 Thermal Analysis

    Science.gov (United States)

    Clayton, J. Louie

    2001-01-01

    This study provides for development and test verification of a thermal model used for prediction of joint heating environments, structural temperatures and seal erosions in the Space Shuttle Reusable Solid Rocket Motor (RSRM) Nozzle Joint-4. The heating environments are a result of rapid pressurization of the joint free volume assuming a leak path has occurred in the filler material used for assembly gap close out. Combustion gases flow along the leak path from nozzle environment to joint O-ring gland resulting in local heating to the metal housing and erosion of seal materials. Analysis of this condition was based on usage of the NASA Joint Pressurization Routine (JPR) for environment determination and the Systems Improved Numerical Differencing Analyzer (SINDA) for structural temperature prediction. Model generated temperatures, pressures and seal erosions are compared to hot fire test data for several different leak path situations. Investigated in the hot fire test program were nozzle joint-4 O-ring erosion sensitivities to leak path width in both open and confined joint geometries. Model predictions were in generally good agreement with the test data for the confined leak path cases. Worst case flight predictions are provided using the test-calibrated model. Analysis issues are discussed based on model calibration procedures.

  12. Critical Review of Phenomenological Models for Studying Transient Combustion of Solid Propellants

    Directory of Open Access Journals (Sweden)

    V.E. Zarko

    2010-06-01

    Full Text Available A brief review of the existing generalizations of phenomenological model of solid propellants combustion, originally formulated by Ya. B. Zel'dovich in 1942, is presented. The model still remains valid and allows the description of different patterns of solid propellant transient combustion. However, over the years after this formulation was published, several publications appeared in the literature with assumptions that are not justified. Shortcomings of different model options and their limits of applicability are discussed. In fact, in some cases the attempts of model generalization are based on rather strong additional assumptions that lead to narrowing the applicability range of the model.

  13. On the oxidation and combustion of AlH{sub 3} a potential fuel for rocket propellants and gas generators

    Energy Technology Data Exchange (ETDEWEB)

    Weiser, Volker; Eisenreich, Norbert; Koleczko, Andreas; Roth, Evelin [Fraunhofer-Institut fuer Chemische Technologie (ICT), Joseph-von-Fraunhoferstrasse 7, 76327 Pfinztal (Germany)

    2007-06-15

    Aluminum hydride is a promising candidate for application in energetic materials and hydrogen storages. E.g. an AP/HTPB rocket propellant filled with alane was calculated for a 100 N s kg{sup -1} higher specific impulse compared to the same concentration of aluminum. Different investigations on {alpha}-AlH{sub 3} polyhedra using thermoanalytical methods and X-ray diffraction were performed to receive a better understanding of dehydration at about 450 K, passivation of the remaining porous aluminum particles and further oxidation. A modeling approach to describe these conversions including diffusion processes, Avrami-Erofeev mechanism and Arrhenius type reaction steps of n-th order were introduced. Results were discussed in comparison to experimental investigations under pressure with model propellants on the base of gelled pure nitromethane and also filled with alane or pure aluminum in concentrations of 5%, 10% and 15%. Both alane and aluminum increase the burning rate on a factor of two correlated with a temperature increase up to 500 K and more. A mesa burning effect at 6 to 10 MPa was indicated by the mixtures with alane. (Abstract Copyright [2007], Wiley Periodicals, Inc.)

  14. Maturation of Structural Health Management Systems for Solid Rocket Motors Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Solid rocket motor cases are subject to a variety of external environmental and loading conditions from cradle-to-grave. These conditions can significantly impact...

  15. Another solid rocket booster is mated with a Boeing Delta II rocket at LP17A, CCAS

    Science.gov (United States)

    1999-01-01

    At Pad 17A, Cape Canaveral Air Station, a fourth and final solid rocket booster (SRB) (right) is moved from the mobile tower by a crane before mating with the Delta II rocket (left). The rocket will be aided by four SRBs to carry the Stardust satellite into space for a close encounter with the comet Wild 2 in January 2004. Using a medium called aerogel, Stardust will capture comet particles flying off the nucleus of the comet, plus collect interstellar dust for later analysis. The collected samples will return to Earth in a Sample Return Capsule to be jettisoned as Stardust swings by Earth in January 2006. Stardust is scheduled to be launched on Feb. 6, 1999.

  16. An Investigation of High-frequency Combustion Oscillations in Liquid-propellant Rocket Engines

    Science.gov (United States)

    Mantler, Raymond L; Tischler, Adelbert O; Massa, Rudolph V

    1953-01-01

    An experimental investigation of high-frequency combustion oscillations (screaming) was conducted with a 100-pound-thrust acid-hydrocarbon rocket engine and a 500-pound-thrust oxygen-fuel rocket engine. The oscillation frequencies could be correlated as a linear function of the parameter C/L, where C is the experimentally measured characteristic velocity and L is the combustion-chamber length. The tendency of the engines to scream increased as chamber length was increased. With engine configurations that normally had a low efficiency, screaming resulted in increased performance; at the same time, a five to tenfold increase in heat-transfer rate occurred. It was possible, however, to achieve good performance without screaming.

  17. Cu–Co–O nano-catalysts as a burn rate modifier for composite solid propellants

    Directory of Open Access Journals (Sweden)

    D. Chaitanya Kumar Rao

    2016-08-01

    Full Text Available Nano-catalysts containing copper–cobalt oxides (Cu–Co–O have been synthesized by the citric acid (CA complexing method. Copper (II nitrate and Cobalt (II nitrate were employed in different molar ratios as the starting reactants to prepare three types of nano-catalysts. Well crystalline nano-catalysts were produced after a period of 3 hours by the calcination of CA–Cu–Co–O precursors at 550 °C. The phase morphologies and crystal composition of synthesized nano-catalysts were examined using Scanning Electron Microscope (SEM, Energy Dispersive Spectroscopy (EDS and Fourier Transform Infrared Spectroscopy (FTIR methods. The particle size of nano-catalysts was observed in the range of 90 nm–200 nm. The prepared nano-catalysts were used to formulate propellant samples of various compositions which showed high reactivity toward the combustion of HTPB/AP-based composite solid propellants. The catalytic effects on the decomposition of propellant samples were found to be significant at higher temperatures. The combustion characteristics of composite solid propellants were significantly improved by the incorporation of nano-catalysts. Out of the three catalysts studied in the present work, CuCo-I was found to be the better catalyst in regard to thermal decomposition and burning nature of composite solid propellants. The improved performance of composite solid propellant can be attributed to the high crystallinity, low agglomeration and lowering the decomposition temperature of oxidizer by the addition of CuCo-I nano-catalyst.

  18. Evaluation of Additives to Reduce Solid Propellant Flammability in Ambient Air.

    Science.gov (United States)

    1975-12-01

    TITLE fund Suhitit*) EVALUATION OF ADDITIVES TO REDUCE SOLID PROPELLANT FLAMMABILITY IN AMBIENT AIR • TYRE OP REPORT A PERIOD COVEREO... BLAST TOKCH WHICH CAN BE PIVOTED TO IMPINGE SPECIMEN (a) Photograph of apparatus. VERTICAL PYREX SHIELD FLAME TORCH SPECIMEN HOLDER

  19. A MEMS-based solid propellant microthruster array for space and military applications

    Science.gov (United States)

    Chaalane, A.; Chemam, R.; Houabes, M.; Yahiaoui, R.; Metatla, A.; Ouari, B.; Metatla, N.; Mahi, D.; Dkhissi, A.; Esteve, D.

    2015-12-01

    Since combustion is an easy way to achieve large quantities of energy from a small volume, we developed a MEMS based solid propellant microthruster array for small spacecraft and micro-air-vehicle applications. A thruster is composed of a fuel chamber layer, a top-side igniter with a micromachined nozzle in the same silicon layer. Layers are assembled by adhesive bonding to give final MEMS array. The thrust force is generated by the combustion of propellant stored in a few millimeter cube chamber. The micro-igniter is a polysilicon resistor deposited on a low stress SiO2/SiNx thin membrane to ensure a good heat transfer to the propellant and thus a low electric power consumption. A large range of thrust force is obtained simply by varying chamber and nozzle geometry parameters in one step of Deep Reactive Ion Etching (DRIE). Experimental tests of ignition and combustion employing home made (DB+x%BP) propellant composed of a DoubleBase and Black-Powder. A temperature of 250 °C, enough to propellant initiation, is reached for 40 mW of electric power. A combustion rate of about 3.4 mm/s is measured for DB+20%BP propellant and thrust ranges between 0.1 and 3,5 mN are obtained for BP ratio between 10% and 30% using a microthruster of 100 μm of throat wide.

  20. Liquid propellant rocket engine combustion simulation with a time-accurate CFD method

    Science.gov (United States)

    Chen, Y. S.; Shang, H. M.; Liaw, Paul; Hutt, J.

    1993-01-01

    Time-accurate computational fluid dynamics (CFD) algorithms are among the basic requirements as an engineering or research tool for realistic simulations of transient combustion phenomena, such as combustion instability, transient start-up, etc., inside the rocket engine combustion chamber. A time-accurate pressure based method is employed in the FDNS code for combustion model development. This is in connection with other program development activities such as spray combustion model development and efficient finite-rate chemistry solution method implementation. In the present study, a second-order time-accurate time-marching scheme is employed. For better spatial resolutions near discontinuities (e.g., shocks, contact discontinuities), a 3rd-order accurate TVD scheme for modeling the convection terms is implemented in the FDNS code. Necessary modification to the predictor/multi-corrector solution algorithm in order to maintain time-accurate wave propagation is also investigated. Benchmark 1-D and multidimensional test cases, which include the classical shock tube wave propagation problems, resonant pipe test case, unsteady flow development of a blast tube test case, and H2/O2 rocket engine chamber combustion start-up transient simulation, etc., are investigated to validate and demonstrate the accuracy and robustness of the present numerical scheme and solution algorithm.

  1. Effect of propellant morphology on acoustics in a planar rocket motor

    Energy Technology Data Exchange (ETDEWEB)

    Daimon, Y.; Jackson, T.L. [University of Illinois at Urbana-Champaign, Center for Simulation of Advanced Rockets, Urbana, IL (United States); Topalian, V. [University of Illinois at Urbana-Champaign, Mechanical Science and Engineering, Urbana, IL (United States); Freund, J.B. [University of Illinois at Urbana-Champaign, Mechanical Science and Engineering, Aerospace Engineering, Urbana, IL (United States); Buckmaster, J. [Buckmaster Research, Urbana, IL (United States)

    2009-03-15

    This paper reports the results of numerical simulations of the acoustics in a two-dimensional (plane) motor using a high-order accurate, low-dissipation numerical solver. For verification we compare solutions to Culick's (AIAA J 4(8):1462-1464, 1966) asymptotic solution for constant injection, and to recent results of Hegab and Kassoy (AIAA J 44(4):812-826, 2006) for a space- and time-dependent mass injection. We present results when the injection boundary condition is described by propellant morphology and by white noise. Morphology strongly affects the amplitude of the longitudinal acoustic modes, and in this connection white noise is not a suitable surrogate. (orig.)

  2. Synergistic and Novel Effects in Composite Solid Propellant Combustion.

    Science.gov (United States)

    1974-10-01

    ac copper chromite . Ferric oxide, Fe?0 , exhibits, as do most k:S ’• • I« Win Mill transition element compounds, a high degree of catalytic... solid phase heat transfer may b^ solved exactly, the solution is V*’ - 1 • (gs - 1) e An overall energy balance yields the adiabatic flame...from the planar solution , but b) the solid phase deviations may be complex. By "simple" it is meant that the de- viation is not oscillatory. Thus, if g

  3. Solid Rocket Booster Large Main and Drogue Parachute Reliability Analysis

    Science.gov (United States)

    Clifford, Courtenay B.; Hengel, John E.

    2009-01-01

    The parachutes on the Space Transportation System (STS) Solid Rocket Booster (SRB) are the means for decelerating the SRB and allowing it to impact the water at a nominal vertical velocity of 75 feet per second. Each SRB has one pilot, one drogue, and three main parachutes. About four minutes after SRB separation, the SRB nose cap is jettisoned, deploying the pilot parachute. The pilot chute then deploys the drogue parachute. The drogue chute provides initial deceleration and proper SRB orientation prior to frustum separation. At frustum separation, the drogue pulls the frustum from the SRB and allows the main parachutes that are mounted in the frustum to unpack and inflate. These chutes are retrieved, inspected, cleaned, repaired as needed, and returned to the flight inventory and reused. Over the course of the Shuttle Program, several improvements have been introduced to the SRB main parachutes. A major change was the replacement of the small (115 ft. diameter) main parachutes with the larger (136 ft. diameter) main parachutes. Other modifications were made to the main parachutes, main parachute support structure, and SRB frustum to eliminate failure mechanisms, improve damage tolerance, and improve deployment and inflation characteristics. This reliability analysis is limited to the examination of the SRB Large Main Parachute (LMP) and drogue parachute failure history to assess the reliability of these chutes. From the inventory analysis, 68 Large Main Parachutes were used in 651 deployments, and 7 chute failures occurred in the 651 deployments. Logistic regression was used to analyze the LMP failure history, and it showed that reliability growth has occurred over the period of use resulting in a current chute reliability of R = .9983. This result was then used to determine the reliability of the 3 LMPs on the SRB, when all must function. There are 29 drogue parachutes that were used in 244 deployments, and no in-flight failures have occurred. Since there are no

  4. Hierarchical Theoretical Methods for Understanding and Predicting Anisotropic Thermal Transport Release in Rocket Propellant Formulations

    Science.gov (United States)

    2016-12-08

    consideration of phase changes including solid-state polymorphism and transitions among solid, liquid, and gas; and tracking of key species...researcher on this project 2016) has begun a tenure- track assistant professor position in the Department of Mechanical Engineering at the University of...left edge of a 100- micron slab was raised to constant temperature at the left edge and insulated on the right edge See Fig. 3. Then the simulations

  5. Cooling process of liquid propellant rocket by means of kerosene-alumina nanofluid

    Directory of Open Access Journals (Sweden)

    Mostafa Mahmoodi

    2016-12-01

    Full Text Available Heat transfer augmentation of kerosene-alumina nanofluid is studied for the possible use in the regenerative cooling channel of semi cryogenic engine. The basic partial differential equations are reduced to ordinary differential equations which are solved using differential transformation method. Velocity and temperature profiles as well as the skin friction coefficient and Nusselt number are determined. The influence of pertinent parameters such as nanofluid volume fraction, viscosity parameter and Eckert number on the flow and heat transfer characteristics is discussed. The results indicate that adding alumina into the fuel of liquid rocket engine (kerosene can be considered as the way of enhancing cooling process of chamber and nozzle walls. Nusselt number is an increasing function of viscosity parameter and nanoparticle volume fraction while it is a decreasing function of Eckert number.

  6. Numerical Study on Similarity of Plume’s Infrared Radiation from Reduced Scaling Solid Rocket

    Directory of Open Access Journals (Sweden)

    Xiaoying Zhang

    2015-01-01

    Full Text Available Similarity of plume radiation between reduced scaling solid rocket models and full scale ones in ground conditions has been taken for investigation. Flow and radiation of plume from solid rockets with scaling ratio from 0.1 to 1 have been computed. The radiative transfer equation (RTE is solved by the finite volume method (FVM in infrared band 2~6 μm. The spectral characteristics of plume gases have been calculated with the weighted-sum-of-gray-gas (WSGG model, and those of the Al2O3 particles have been solved by the Mie scattering model. Our research shows that, with the decreasing scaling ratio of the rocket engine, the radiation intensity of the plume decreases with 1.5~2.5 power of the scaling ratio. The infrared radiation of the plume gases shows a strong spectral dependency, while that of the Al2O3 particles shows grey property. Spectral radiation intensity of the high temperature core of the solid rocket plume increases greatly in the peak absorption spectrum of plume gases. Al2O3 particle is the major radiation composition in the rocket plume, whose scattering coefficient is much larger than its absorption coefficient. There is good similarity between spectral variations of plumes from different scaling solid rockets. The directional plume radiation rises with the increasing azimuth angle.

  7. Eulerian multi-fluid models for the simulation of dynamics and coalescence of particles in solid propellant combustion

    Science.gov (United States)

    Doisneau, F.; Laurent, F.; Murrone, A.; Dupays, J.; Massot, M.

    2013-02-01

    The accurate simulation of polydisperse sprays undergoing coalescence in unsteady gaseous flows is a crucial issue. In solid rocket motors, the internal flow depends strongly on the alumina droplet size distribution, which spreads up with coalescence. Yet solving for unsteady two-phase flows with high accuracy on the droplet sizes is a challenge for both modeling and scientific computing. As an alternative to Lagrangian approaches, a wide range of Eulerian models have been recently developed to describe the disperse liquid phase at a lower cost, with an easier coupling to the gaseous phase and with massively parallel codes. Among these models, the multi-fluid model allows the detailed description of polydispersity and size/velocity correlations by separately solving fluids of size-sorted droplets, the so-called sections. The existing one size moment method, which describes the size distribution with one size moment per section, provides simple and fast resolution for coalescence. On the other hand, a two size moment method has been suggested to reduce the number of sections but it lacks an efficient coalescence resolution method. After introducing a new strategy for two size moment coalescence, the two methods are compared on various configurations in a research code and an industrial-oriented code, in order to conclude on computational accuracy and cost. Then the paper aims at describing the most efficient approach for multi-dimensional unsteady and eventually coalescing rocket chamber simulations. Its objective is threefold: first, to validate the two size moment method by comparing simulations to reference solutions and dedicated experimental measurements conducted at ONERA, second to study the efficiency and robustness of both methods, third, to draw some firm conclusions about the necessity to use the one size moment or two size moment method to simulate solid propellant alumina sprays. We finally perform the first simulations of coalescence in realistic 2D

  8. Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical Analyzer

    Science.gov (United States)

    2017-03-29

    Article 3. DATES COVERED (From - To) 04 August 2016 – 29 March 2017 4. TITLE AND SUBTITLE Measuring Propellant Stress Relaxation Modulus Using Dynamic...ERC 14. ABSTRACT A method for determining the stress relaxation master curve of solid rocket propellants was developed. The propellant was tested in... stress relaxation modulus, Mpa; E∞ = long-term equilibrium modulus, Mpa; E0 = storage modulus, Mpa; E0 0 = loss modulus, Mpa; F = Fisher test

  9. Workshop Report: Fundamental Reactions in Solid Propellant Combustion

    Science.gov (United States)

    1979-05-01

    combustion conditions. 6. What effect might a pressure-induced phase transition to a polymorph other than 6- HMX have on the pressure slope break during...pure HMX as well. Nevertheless, it is recommended that the high pressure polymorphs of HMX and RDX be determined. It was also felt that there...plateau burning phenomena E. Solid phase, surface, gas phase reactions F. Phase transitions: melting, vaporization, polymorphs G. Flame

  10. Treatment of Solid Rocket Motors that Complies with Established Protocols to Ensure Planetary Protection

    Science.gov (United States)

    Stefanski, Philip L.; Soler-Luna, Adrian

    2017-01-01

    This presentation discusses recent work being conducted by the National Aeronautics and Space Administration (NASA) at Marshall Space Flight Center (MSFC) to evaluate various methods that could be employed to provide for planetary protection of those solar system bodies that are candidates for extraterrestrial life, thus preventing contamination of such bodies. MSFC is presently involved in the development phase of the Europa Lander De-Orbital Stage (DOS) braking motor. In order to prevent bio-contamination of this Jovian satellite, three paths are currently being considered. The first is (1) Bio-Reduction of those microscopic organisms in or on the vehicle (in this case a solid rocket motor (SRM)) that might otherwise be transported during the mission. Possible methods being investigated include heat sterilization, application or incorporation of biocide materials, and irradiation. While each method can be made to work, effects on the SRM's components (propellant, liner, insulation, etc.) could well prove deleterious. A second path would be use of (2) Bio-Barrier material(s). So long as such barrier(s) can maintain their integrity, planetary protection should be afforded. Under the harsh conditions encountered during extended spaceflight (vacuum, temperature extremes, radiation), however, such barrier(s) could well experience a breach. Finally, a third path would be to perform (3) Pyrotechnic Sterilization of the SRM during its end-of-mission phase. Multiple pyrotechnic units would be triggered to ensure activation of such an event and provide for a final sterilization before vehicle impact. In light of Europa's stringent bio-reduction targets, the final and best choice to minimize risk will probably be some combination of the above.

  11. An Internal Thermal Environment Model of an Aluminized Solid Rocket Motor with Experimental Validation

    Science.gov (United States)

    Martin, Heath T.

    2015-01-01

    Due to the severity of the internal solid rocket motor (SRM) environment, very few direct measurements of that environment exist; therefore, the appearance of such data provides a unique opportunity to assess current thermal/fluid modeling capabilities. As part of a previous study of SRM internal insulation performance, the internal thermal environment of a laboratory-scale SRM featuring aluminized propellant was characterized with two types of custom heat-flux calorimeters: one that measured the total heat flux to a graphite slab within the SRM chamber and another that measured the thermal radiation flux. Therefore, in the current study, a thermal/fluid model of this lab-scale SRM was constructed using ANSYS Fluent to predict not only the flow field structure within the SRM and the convective heat transfer to the interior walls, but also the resulting dispersion of alumina droplets and the radiative heat transfer to the interior walls. The dispersion of alumina droplets within the SRM chamber was determined by employing the Lagrangian discrete phase model that was fully coupled to the Eulerian gas-phase flow. The P1-approximation was engaged to model the radiative heat transfer through the SRM chamber where the radiative contributions of the gas phase were ignored and the aggregate radiative properties of the alumina dispersion were computed from the radiative properties of its individual constituent droplets, which were sourced from literature. The convective and radiative heat fluxes computed from the thermal/fluid model were then compared with those measured in the lab-scale SRM test firings and the modeling approach evaluated.

  12. Uncertainty Quantification of Non-linear Oscillation Triggering in a Multi-injector Liquid-propellant Rocket Combustion Chamber

    Science.gov (United States)

    Popov, Pavel; Sideris, Athanasios; Sirignano, William

    2014-11-01

    We examine the non-linear dynamics of the transverse modes of combustion-driven acoustic instability in a liquid-propellant rocket engine. Triggering can occur, whereby small perturbations from mean conditions decay, while larger disturbances grow to a limit-cycle of amplitude that may compare to the mean pressure. For a deterministic perturbation, the system is also deterministic, computed by coupled finite-volume solvers at low computational cost for a single realization. The randomness of the triggering disturbance is captured by treating the injector flow rates, local pressure disturbances, and sudden acceleration of the entire combustion chamber as random variables. The combustor chamber with its many sub-fields resulting from many injector ports may be viewed as a multi-scale complex system wherein the developing acoustic oscillation is the emergent structure. Numerical simulation of the resulting stochastic PDE system is performed using the polynomial chaos expansion method. The overall probability of unstable growth is assessed in different regions of the parameter space. We address, in particular, the seven-injector, rectangular Purdue University experimental combustion chamber. In addition to the novel geometry, new features include disturbances caused by engine acceleration and unsteady thruster nozzle flow.

  13. Aluminum Agglomeration and Trajectory in Solid Rocket Motors

    Science.gov (United States)

    2007-08-30

    agglomerates coming off the surface of a burning propellant. These methods include film studies 6, quench particles combustion bomb (QPCB) devices similar...34the stepwise oxidation of aluminum (that) is caused by the sequence of polymorphic phase transitions occurring in the growing oxide film ",2 5 . 25... cinematography data from China Lake. Task 2.2, Aluminum Agglomeration Model Selection (SEA/BYU/ATK Task) Part of the model selection task has already been

  14. An Investigation of Particulate Behavior in Solid Rocket Motors

    Science.gov (United States)

    1981-06-01

    together with corresponding pictures taken with the high- speed cinematography method. The latter were obtained from the films taken by Diloreto [Ref...High-Speed Cinematography Metallized Propellants Holography Light Scattering 20. AT~tFT (C.U50us on Fever aide if ft oGo. IF d Ids t I- I IO 1101...Page I. INTRODUCTION ---------------------------------- 10 II. -METHOD OF INVESTIGATION ----------------------- 19 III. HIGH-SPEED CINEMATOGRAPHY

  15. Design criteria of launching rockets for burst aerial shells

    Energy Technology Data Exchange (ETDEWEB)

    Kuwahara, T.; Takishita, Y.; Onda, T.; Shibamoto, H.; Hosaya, F. [Hosaya Kako Co. Ltd (Japan); Kubota, N. [Mitsubishi Electric Corporation (Japan)

    2000-04-01

    Rocket motors attached to large-sized aerial shells are proposed to compensate for the increase in the lifting charge in the mortar and the thickness of the shell wall. The proposal is the result of an evaluation of the performance of solid propellants to provide information useful in designing launch rockets for large-size shells. The propellants composed of ammonium perchlorate and hydroxy-terminated polybutadiene were used to evaluate the ballistic characteristics such as the relationship between propellant mass and trajectories of shells and launch rockets. In order to obtain an optimum rocket design, the evaluation also included a study of the velocity and height of the rocket motor and shell separation. A launch rocket with a large-sized shell (84.5 cm in diameter) was designed to verify the effectiveness of this class of launch system. 2 refs., 6 figs.

  16. ROCKET

    OpenAIRE

    CRUZ BARCELÓ, JULIA

    2015-01-01

    [In] The main objective of this project was to perform a stop-motion animation short film with my colleague Marta Soriano Gil, student of Fine Arts Degree at the Polytechnic University of Valencia, like myself. This project collects a year’s work developing the film, from the origins of the idea and the early stages of production to the final editing and audiovisual document. Rocket, is a project that has encompassed many disciplines, from the initial draws, designs and conc...

  17. Hydromining a full-scale class 1.1 solid rocket motor

    Science.gov (United States)

    Young, M. F.; Chleborad, O. W.; Cova, P. J.; Eccli, D. B.; Glad, T. J.; Kunkle, D. M.

    1993-11-01

    A class 1.1 full-scale Stage 2 Small Intercontinental Ballistic Missile (SICBM) was successfully deloaded using high pressure water jets at approximately 689 N (10,000 psia). The SICBM Stage 2 rocket motor contains approximately 2948 kg (6500 lb) of Class 1.1 propellant (nitrato ester polyester based propellant). A vertical hydromining facility was constructed to simulate a production operational mode in which the system water was continuously recycled to the high pressure water jet nozzles. A series of four in-line carbon absorption filter pairs were used to reduce nitrato ester concentrations to acceptable safety levels (e.g., less than 1 ppm of nitroglycerin). This paper will summarize the operating conditions used and the results of the successful study.

  18. Trinitromethyl Heterocyclic Oxidizers as a Solid Propellant Ingredient Final Report CRADA No TC02146.0

    Energy Technology Data Exchange (ETDEWEB)

    Pagoria, P. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States); Racoveanu, A. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States)

    2017-08-15

    This was a collaborative effort between Lawrence Livermore National Security, LLC as manager and operator of Lawrence Livermore National Laboratory (LLNL) and Physical Sciences, Inc. (PSI), to develop a synthesis of two novel energetic heterocyclic oxidizers as possible replacements for ammonium perchlorate (AP) in rocket propellant formulations. This CRADA resulted from the award of the Phase I Small Business Technology Transfer (STTR) from DOD. The CRADA consisted of two phases. The goal for Phase 1 was to produce a new oxidizer called TNMDNP. Phase 2 is optional (based on the success of Phase 1) and the goal of Phase 2 (optional) was to produce a new oxidizer called TNMDNT. Phase 2 tasks would be performed based on the successful results of Phase 1.

  19. The Use of Ion Vapor Deposited Aluminum (IVD) for the Space Shuttle Solid Rocket Booster (SRB)

    Science.gov (United States)

    Novak, Howard L.

    2003-01-01

    This viewgraph representation provides an overview of the use of ion vapor deposited aluminum (IVD) for use in the Space Shuttle Solid Rocket Booster (SRB). Topics considered include: schematics of ion vapor deposition system, production of ion vapor deposition system, IVD vs. cadmium coated drogue ratchets, corrosion exposure facilities and tests, seawater immersion facilities and tests and continued research and development issues.

  20. Preliminary analysis of hybrid rockets for launching nanosats into LEO

    OpenAIRE

    Costa, Fernando de Souza; Vieira, Ricardo

    2010-01-01

    This work determines the preliminary mass distribution of hybrid rockets using 98% H2O2 and solid paraffin mixed with aluminum as propellants. An iterative process is used tocalculate the rocket performance characteristics and to determine the inert mass fractionfrom given initial conditions. It is considered a mission to place a 20 kg payload into a 300 km circular equatorial orbit by air launched and ground launched hybrid rockets usingthree stages. The results indicate total initial masses...

  1. Structural Assessment of Solid Propellant Grains (l’Evaluation structurale des blocs de poudre a’ propergol solide)

    Science.gov (United States)

    1997-12-01

    Life of Solid Propellant Systems, 6-10 May 1996, Athens. [8] Chevalier, M.; Laurent, J.M.; Programme General d’Essais Vieillissement . Application a...analysis, the finite 2-4 element solution procedure and interpretation of results, and the basic material characterisation inputs and the definition of... DEFINITION 3-2 3.2.1 Problem Definition 3-2 3.2.1.1 Two Dimensional Analysis 3-4 3.2.1.2 Three Dimensional Analysis 3-4 3.2.2 Finite Element Model

  2. Optimization of Construction of the rocket-assisted projectile

    Directory of Open Access Journals (Sweden)

    Arkhipov Vladimir

    2017-01-01

    Full Text Available New scheme of the rocket motor of rocket-assisted projectile providing the increase in distance of flight due to controlled and optimal delay time of ignition of the solid-propellant charge of the SRM and increase in reliability of initiation of the SRM by means of the autonomous system of ignition excluding the influence of high pressure gases of the propellant charge in the gun barrel has been considered. Results of the analysis of effectiveness of using of the ignition delay device on motion characteristics of the rocket-assisted projectile has been presented.

  3. Failure mode and effects analysis (FMEA) for the Space Shuttle solid rocket motor

    Science.gov (United States)

    Russell, D. L.; Blacklock, K.; Langhenry, M. T.

    1988-01-01

    The recertification of the Space Shuttle Solid Rocket Booster (SRB) and Solid Rocket Motor (SRM) has included an extensive rewriting of the Failure Mode and Effects Analysis (FMEA) and Critical Items List (CIL). The evolution of the groundrules and methodology used in the analysis is discussed and compared to standard FMEA techniques. Especially highlighted are aspects of the FMEA/CIL which are unique to the analysis of an SRM. The criticality category definitions are presented and the rationale for assigning criticality is presented. The various data required by the CIL and contribution of this data to the retention rationale is also presented. As an example, the FMEA and CIL for the SRM nozzle assembly is discussed in detail. This highlights some of the difficulties associated with the analysis of a system with the unique mission requirements of the Space Shuttle.

  4. Ignition transient modelling for the Space Shuttle Advanced Solid Rocket Motor

    Science.gov (United States)

    Eagar, M. A.; Luke, G. D.; Stockham, L. W.

    1993-06-01

    Prediction of the ignition transient for the Advanced Solid Rocket Motor (ASRM) for the Space Shuttle presents an unusual set of modelling challenges because of its high length-to-diameter ratio and complex internal flow environment. A review of ignition modelling experience on the Shuttle Redesigned Solid Rocket Motor (RSRM), which is similar in size and configuration to the ASRM, reveals that classical igniter design theory and modelling methods under-predict, by a factor of two, the measured pressure and thrust rise rates experienced on the RSRM. This paper (1) reviews the Titan and Shuttle SRM test experience, (2) presents the results of 0-Dimensional (0-D) and 1-Dimensional (1-D) analysis of the RSRM and ASRM motors, and (3) addresses the need for advanced analysis techniques, as they relate to ASRM ignition transient modelling requirements and igniter design drivers.

  5. Test and Analysis of Solid Rocket Motor Nozzle Ablative Materials

    Science.gov (United States)

    Clayton, J. Louie

    2017-01-01

    Asbestos free solid motor internal insulation samples were tested at the MSFC Hyperthermal Facility. Objectives of the test were to gather data for analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Instrumentation included use of total calorimeters, thermocouples, and a surface pyrometer for surface temperature measurement. Post-test sample forensics involved measurement of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero-thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

  6. Results of Small-scale Solid Rocket Combustion Simulator testing at Marshall Space Flight Center

    Science.gov (United States)

    Goldberg, Benjamin E.; Cook, Jerry

    1993-01-01

    The Small-scale Solid Rocket Combustion Simulator (SSRCS) program was established at the Marshall Space Flight Center (MSFC), and used a government/industry team consisting of Hercules Aerospace Corporation, Aerotherm Corporation, United Technology Chemical Systems Division, Thiokol Corporation and MSFC personnel to study the feasibility of simulating the combustion species, temperatures and flow fields of a conventional solid rocket motor (SRM) with a versatile simulator system. The SSRCS design is based on hybrid rocket motor principles. The simulator uses a solid fuel and a gaseous oxidizer. Verification of the feasibility of a SSRCS system as a test bed was completed using flow field and system analyses, as well as empirical test data. A total of 27 hot firings of a subscale SSRCS motor were conducted at MSFC. Testing of the Small-scale SSRCS program was completed in October 1992. This paper, a compilation of reports from the above team members and additional analysis of the instrumentation results, will discuss the final results of the analyses and test programs.

  7. Real-Time X-ray Radiography Diagnostics of Components in Solid Rocket Motors

    Science.gov (United States)

    Cortopassi, A. C.; Martin, H. T.; Boyer, E.; Kuo, K. K.

    2012-01-01

    Solid rocket motors (SRMs) typically use nozzle materials which are required to maintain their shape as well as insulate the underlying support structure during the motor operation. In addition, SRMs need internal insulation materials to protect the motor case from the harsh environment resulting from the combustion of solid propellant. In the nozzle, typical materials consist of high density graphite, carbon-carbon composites and carbon phenolic composites. Internal insulation of the motor cases is typically a composite material with carbon, asbestos, Kevlar, or silica fibers in an ablative matrix such as EPDM or NBR. For both nozzle and internal insulation materials, the charring process occurs when the hot combustion products heat the material intensely. The pyrolysis of the matrix material takes away a portion of the thermal energy near the wall surface and leaves behind a char layer. The fiber reinforcement retains the porous char layer which provides continued thermal protection from the hot combustion products. It is of great interest to characterize both the total erosion rates of the material and the char layer thickness. By better understanding of the erosion process for a particular ablative material in a specific flow environment, the required insulation material thickness can be properly selected. The recession rates of internal insulation and nozzle materials of SRMs are typically determined by testing in some sort of simulated environment; either arc-jet testing, flame torch testing, or subscale SRMs of different size. Material recession rates are deduced by comparison of pre- and post-test measurements and then averaging over the duration of the test. However, these averaging techniques cannot be used to determine the instantaneous recession rates of the material. Knowledge of the variation in recession rates in response to the instantaneous flow conditions during the motor operation is of great importance. For example, in many SRM configurations

  8. Diffusion of plasticizer in a solid propellant based on hydroxyl-terminated polybutadiene

    Directory of Open Access Journals (Sweden)

    Juliano Libardi

    2010-01-01

    Full Text Available The aim of this work was to determine the dioctyl azelate (DOZ plasticizer diffusion coefficient (D for samples containing the interfaces of rubber, liner and solid composite propellant based on hydroxyl-terminated polybutadiene (HTPB. The samples used in the diffusion study were aged up to 31 days after the cure period at 80 °C. A computer program implementing a mathematical model of Fick's second Law of diffusion was developed to calculate the diffusion coefficient based on concentration data obtained from gas chromatographic analyses. The effects of the diffusion phenomenon were also investigated by Shore A hardness and scanning electron microscope (SEM techniques. These analyses were carried out using samples aged at room temperature and at 80 °C. The hardness results showed an increasing trend for the samples aged at room temperature; however in the tests carried out at 80 °C they showed the opposite trend. The SEM analyses detected meaningful changes in the surface morphology of the propellant for both aging temperatures.

  9. Modeling and simulation of the debonding process of composite solid propellants

    Science.gov (United States)

    Feng, Tao; Xu, Jin-sheng; Han, Long; Chen, Xiong

    2017-07-01

    In order to study the damage evolution law of composite solid propellants, the molecular dynamics particle filled algorithm was used to establish the mesoscopic structure model of HTPB(Hydroxyl-terminated polybutadiene) propellants. The cohesive element method was employed for the adhesion interface between AP(Ammonium perchlorate) particle and HTPB matrix and the bilinear cohesive zone model was used to describe the mechanical response of the interface elements. The inversion analysis method based on Hooke-Jeeves optimization algorithm was employed to identify the parameters of cohesive zone model(CZM) of the particle/binder interface. Then, the optimized parameters were applied to the commercial finite element software ABAQUS to simulate the damage evolution process for AP particle and HTPB matrix, including the initiation, development, gathering and macroscopic crack. Finally, the stress-strain simulation curve was compared with the experiment curves. The result shows that the bilinear cohesive zone model can accurately describe the debonding and fracture process between the AP particles and HTPB matrix under the uniaxial tension loading.

  10. Ignition by laser radiation and combustion of composite solid propellants with bimetal powders

    Science.gov (United States)

    Korotkikh, A. G.; Arkhipov, V. A.; Glotov, O. G.; Zolotorev, N. N.

    2017-05-01

    The use of metal powder (usually aluminum) as a fuel in composite solid propellants (CSPs) for propulsion is the most energy efficient method that allows improvement of combustion characteristics of propellants in the combustion chamber and specific impulse. This paper presents the experimental data of the ignition and combustion processes of CSPs containing Alex aluminum nanopowder and mixtures of Alex/Fe and Alex/B nanopowders. It was found that the introduction of Alex/Fe in CSPs leads to 1.3-1.9 times decrease in the ignition time at q = 55-220 W/cm2 and to 1.3-1.4 times increase in the burning rate at p = 2.2-7.5 MPa with respect to that for basic CSP with Alex. When introducing Alex/B in CSP, the ignition times are 1.2-1.4 fold decreased, and the burning rate is practically unchanged. However, the agglomeration is significantly enhanced, which is manifested through the increase in the agglomerate particles content in condensed combustion products by a factor of 1.8-2.2, at 1.6-1.7 fold increase of the agglomerates mean diameter for CSP with Alex/B.

  11. Condensed-Phase Processes during Solid Propellant Combustion. 1. Preliminary Chemical and Microscopic Examination of Extinguished Propellant Samples

    Science.gov (United States)

    1990-06-01

    to nitramines or nitramine propellants, although in one studyC the surface layers of a quenched RDX-polyester composition were extracted with benzene...performance liquid chromatography (HPLC). In some cases scraped and unscraped burned surfaces were exmined by photoacoustic Fourier-transform infrared...formed, relative to RDX and HMX. It is thus estimated that the nitrosamines are present in amounts as high as 1-10% of the unreacted RDX in some cases

  12. Response of selected plant and insect species to simulated solid rocket exhaust mixtures and to exhaust components from solid rocket fuels

    Science.gov (United States)

    Heck, W. W.; Knott, W. M.; Stahel, E. P.; Ambrose, J. T.; Mccrimmon, J. N.; Engle, M.; Romanow, L. A.; Sawyer, A. G.; Tyson, J. D.

    1980-01-01

    The effects of solid rocket fuel (SRF) exhaust on selected plant and and insect species in the Merritt Island, Florida area was investigated in order to determine if the exhaust clouds generated by shuttle launches would adversely affect the native, plants of the Merritt Island Wildlife Refuge, the citrus production, or the beekeeping industry of the island. Conditions were simulated in greenhouse exposure chambers and field chambers constructed to model the ideal continuous stirred tank reactor. A plant exposure system was developed for dispensing and monitoring the two major chemicals in SRF exhaust, HCl and Al203, and for dispensing and monitoring SRF exhaust (controlled fuel burns). Plants native to Merritt Island, Florida were grown and used as test species. Dose-response relationships were determined for short term exposure of selected plant species to HCl, Al203, and mixtures of the two to SRF exhaust.

  13. Thermo-Structural Response Caused by Structure Gap and Gap Design for Solid Rocket Motor Nozzles

    Directory of Open Access Journals (Sweden)

    Lin Sun

    2016-06-01

    Full Text Available The thermo-structural response of solid rocket motor nozzles is widely investigated in the design of modern rockets, and many factors related to the material properties have been considered. However, little work has been done to evaluate the effects of structure gaps on the generation of flame leaks. In this paper, a numerical simulation was performed by the finite element method to study the thermo-structural response of a typical nozzle with consideration of the structure gap. Initial boundary conditions for thermo-structural simulation were defined by a quasi-1D model, and then coupled simulations of different gap size matching modes were conducted. It was found that frictional interface treatment could efficiently reduce the stress level. Based on the defined flame leak criteria, gap size optimization was carried out, and the best gap matching mode was determined for designing the nozzle. Testing experiment indicated that the simulation results from the proposed method agreed well with the experimental results. It is believed that the simulation method is effective for investigating thermo-structural responses, as well as designing proper gaps for solid rocket motor nozzles.

  14. Performance of reinforced polymer ablators exposed to a solid rocket motor exhaust. Technical report

    Energy Technology Data Exchange (ETDEWEB)

    Boyer, C.; Burgess, T.; Bowen, J.; Deloach, K.; Talmy, I.

    1992-10-01

    Summarized in this report is the effort by the Naval Surface Warfare Center Dahlgren Division (NSWCDD) and FMC Corporation (a launcher manufacturer) to identify new high performance ablators suitable for use on Navy guided missile launchers (GML) and ships' structures. The goal is to reduce ablator erosion by 25 to 50 percent compared to that of the existing ablators such as MXBE350 (rubbermodified phenolic containing glass fiber reinforcement). This reduction in erosion would significantly increase the number of new missiles with higher-thrust, longer burn rocket motors that can be launched prior to ablator refurbishment. In fact, there are a number of new Navy missiles being considered for development and introduction into existing GML: e.g., the Antisatellite Missile (ASM) and the Theater High-Altitude Area Defense (THAAD) Missile. The U.S. Navy experimentally evaluated the eight best fiber-reinforced, polymer composites from a possible field of 25 off-the-shelf ablators previously screened by FMC Corporation. They were tested by the Navy in highly aluminized solid rocket motor exhaust plumes to determine their ability to resist erosion and to insulate.... Ablator, Guided Missile Launchers, Erosion, Tactical missiles, Convective heating, Solid rocket motors, Aluminum oxide particles.

  15. A Computer Program for the Prediction of Solid Propellant Rocket Motor Performance. Volume 3

    Science.gov (United States)

    1975-07-01

    8217»^»«">»f)B.»ui» !~»^«*.«1er-" ,*if’ * I*/*»1«•>(-»-*■%•* •j"~-*-n*v»tw-«*-,»4 4-ei-. a» ^«.^^-t^^rfV^tit-^^-j»^ -rf «.i *’•. P. fkk . k.r-*kP» »■•-.« • T

  16. Analysis of large solid propellant rocket engine exhaust plumes using the direct simulation Monte Carlo method

    Science.gov (United States)

    Hueser, J. E.; Brock, F. J.; Melfi, L. T., Jr.; Bird, G. A.

    1984-01-01

    A new solution procedure has been developed to analyze the flowfield properties in the vicinity of the Inertial Upper Stage/Spacecraft during the 1st stage (SRMI) burn. Continuum methods are used to compute the nozzle flow and the exhaust plume flowfield as far as the boundary where the breakdown of translational equilibrium leaves these methods invalid. The Direct Simulation Monte Carlo (DSMC) method is applied everywhere beyond this breakdown boundary. The flowfield distributions of density, velocity, temperature, relative abundance, surface flux density, and pressure are discussed for each species for 2 sets of boundary conditions: vacuum and freestream. The interaction of the exhaust plume and the freestream with the spacecraft and the 2-stream direct interaction are discussed. The results show that the low density, high velocity, counter flowing free-stream substantially modifies the flowfield properties and the flux density incident on the spacecraft. A freestream bow shock is observed in the data, located forward of the high density region of the exhaust plume into which the freestream gas does not penetrate. The total flux density incident on the spacecraft, integrated over the SRM1 burn interval is estimated to be of the order of 10 to the 22nd per sq m (about 1000 atomic layers).

  17. Reusable Solid Rocket Motor - Accomplishment, Lessons, and a Culture of Success

    Science.gov (United States)

    Moore, D. R.; Phelps, W. J.

    2011-01-01

    The Reusable Solid Rocket Motor (RSRM) represents the largest solid rocket motor (SRM) ever flown and the only human-rated solid motor. High reliability of the RSRM has been the result of challenges addressed and lessons learned. Advancements have resulted by applying attention to process control, testing, and postflight through timely and thorough communication in dealing with all issues. A structured and disciplined approach was taken to identify and disposition all concerns. Careful consideration and application of alternate opinions was embraced. Focus was placed on process control, ground test programs, and postflight assessment. Process control is mandatory for an SRM, because an acceptance test of the delivered product is not feasible. The RSRM maintained both full-scale and subscale test articles, which enabled continuous improvement of design and evaluation of process control and material behavior. Additionally RSRM reliability was achieved through attention to detail in post flight assessment to observe any shift in performance. The postflight analysis and inspections provided invaluable reliability data as it enables observation of actual flight performance, most of which would not be available if the motors were not recovered. RSRM reusability offered unique opportunities to learn about the hardware. NASA is moving forward with the Space Launch System that incorporates propulsion systems that takes advantage of the heritage Shuttle and Ares solid motor programs. These unique challenges, features of the RSRM, materials and manufacturing issues, and design improvements will be discussed in the paper.

  18. Flight demonstration of flight termination system and solid rocket motor ignition using semiconductor laser initiated ordnance

    Science.gov (United States)

    Schulze, Norman R.; Maxfield, B.; Boucher, C.

    1995-01-01

    Solid State Laser Initiated Ordnance (LIO) offers new technology having potential for enhanced safety, reduced costs, and improved operational efficiency. Concerns over the absence of programmatic applications of the technology, which has prevented acceptance by flight programs, should be abated since LIO has now been operationally implemented by the Laser Initiated Ordnance Sounding Rocket Demonstration (LOSRD) Program. The first launch of solid state laser diode LIO at the NASA Wallops Flight Facility (WFF) occurred on March 15, 1995 with all mission objectives accomplished. This project, Phase 3 of a series of three NASA Headquarters LIO demonstration initiatives, accomplished its objective by the flight of a dedicated, all-LIO sounding rocket mission using a two-stage Nike-Orion launch vehicle. LIO flight hardware, made by The Ensign-Bickford Company under NASA's first Cooperative Agreement with Profit Making Organizations, safely initiated three demanding pyrotechnic sequence events, namely, solid rocket motor ignition from the ground and in flight, and flight termination, i.e., as a Flight Termination System (FTS). A flight LIO system was designed, built, tested, and flown to support the objectives of quickly and inexpensively putting LIO through ground and flight operational paces. The hardware was fully qualified for this mission, including component testing as well as a full-scale system test. The launch accomplished all mission objectives in less than 11 months from proposal receipt. This paper concentrates on accomplishments of the ordnance aspects of the program and on the program's implementation and results. While this program does not generically qualify LIO for all applications, it demonstrated the safety, technical, and operational feasibility of those two most demanding applications, using an all solid state safe and arm system in critical flight applications.

  19. Oxidation/Biodegradation of Solid Propellants Used in Legacy Chemical Rounds

    National Research Council Canada - National Science Library

    Guelta, Mark A; Beck, Andrew S

    2007-01-01

    .... With the planned destruction of the U.S. chemical agent inventory, the associated propellant charges and the now antiquated propellants in storage for use in high energy rounds are awaiting disposal...

  20. Control of the self-oscillation of amplitude of vibration combustion in a liquid-propellant rocket engine by solving the system of equations that describe this regime of combustion

    Science.gov (United States)

    Gotsulenko, V. V.

    2010-07-01

    Periodic solutions of the degenerate system of equations of nonstationary motion of a medium in a liquid-propellant rocket engine were obtained, with the aid of which the possibility of lowering the amplitude of the longitudinal self-oscillations of vibration combustion or their complete removal has been substantiated.

  1. Self-oscillations of an unstable fuel combustion in the combustion chamber of a liquid-propellant rocket engine

    Science.gov (United States)

    Gotsulenko, V. V.; Gotsulenko, V. N.

    2013-01-01

    The form of the self-oscillations of a vibrating combustion of a fuel in the combustion chamber of a liquidpropellant rocket engine, caused by the fuel-combustion lag and the heat release, was determined. The character of change in these self-oscillations with increase in the time of the fuel-combustion lag was investigated.

  2. Evaluation of Flygt Propeller Xixers for Double Shell Tank (DST) High Level Waste Auxiliary Solids Mobilization

    Energy Technology Data Exchange (ETDEWEB)

    PACQUET, E.A.

    2000-07-20

    The River Protection Project (RPP) is planning to retrieve radioactive waste from the single-shell tanks (SST) and double-shell tanks (DST) underground at the Hanford Site. This waste will then be transferred to a waste treatment plant to be immobilized (vitrified) in a stable glass form. Over the years, the waste solids in many of the tanks have settled to form a layer of sludge at the bottom. The thickness of the sludge layer varies from tank to tank, from no sludge or a few inches of sludge to about 15 ft of sludge. The purpose of this technology and engineering case study is to evaluate the Flygt{trademark} submersible propeller mixer as a potential technology for auxiliary mobilization of DST HLW solids. Considering the usage and development to date by other sites in the development of this technology, this study also has the objective of expanding the knowledge base of the Flygt{trademark} mixer concept with the broader perspective of Hanford Site tank waste retrieval. More specifically, the objectives of this study delineated from the work plan are described.

  3. Wind tunnel investigation of Space Shuttle Solid Rocket Booster drogue parachutes and deployment concepts

    Science.gov (United States)

    Bacchus, D. L.; Vickers, J. R.; Foughner, J. T., Jr.

    1975-01-01

    A wind tunnel test has been conducted on one-eighth scale models of the Space Shuttle Solid Rocket Booster drogue parachute system. The test included an investigation of four candidate drogue deployment concepts and a parametric steady state drag study of 20-degree conical ribbon parachutes. The results show that at least two of the four deployment concepts tested are viable candidates for the full scale deployment system. The interference free steady state drag results obtained show excellent agreement with available drop test results on large 20-degree conical ribbon parachutes.

  4. Drogue parachute deployment dynamics of the space shuttle solid rocket booster

    Science.gov (United States)

    Banerjee, A. K.; Utreja, L. R.

    1975-01-01

    Theoretical analysis and experimental investigations are presented for two deployment concepts of the drogue parachute deployment for the space shuttle solid rocket booster. The analysis represents the motion of three coupled rigid bodies, with one of the bodies having variable mass and moment of inertia. The unfurling process of the parachute from the bag is idealized as the flow of a continuum out of a control surface. The pilot parachute or tethered nose cap is modeled as a flexible pendulum with a nonlinear spring and a moving support. Measured wind tunnel test data compare reasonably well with the theory.

  5. Hydrochloric acid aerosol and gaseous hydrogen chloride partitioning in a cloud contaminated by solid rocket exhaust

    Science.gov (United States)

    Sebacher, D. I.; Bendura, R. J.; Wornom, D. E.

    1980-01-01

    Partitioning of hydrogen chloride between hydrochloric acid aerosol and gaseous HCl in the lower atmosphere was experimentally investigated in a solid rocket exhaust cloud diluted with humid ambient air. Airborne measurements were obtained of gaseous HCl, total HCl, relative humidity and temperature to evaluate the conditions under which aerosol formation occurs in the troposphere in the presence of hygroscopic HCl vapor. Equilibrium predictions of HCl aerosol formation accurately predict the measured HCl partitioning over a range of total HCl concentrations from 0.6 to 16 ppm.

  6. Comparative Study on Options for High-Speed Intercontinental Passenger Transports: Air-Breathing- vs. Rocket-Propelled

    OpenAIRE

    Sippel, Martin; Klevanski, Josef; Steelant, Johan

    2005-01-01

    This paper investigates the technical options for high-speed intercontinental passenger transports on a preliminary basis. Horizontal take-off hypersonic air-breathing airliners are assessed as well as vertical take-off, rocket powered stages, capable of a safe atmospheric reentry. The study includes a preliminary sizing and performance assessment of all investigated vehicles and compares characteristic technical and passenger environment data. The aerodynamic shape of the air-bre...

  7. SRM propellant and polymer materials structural test program

    Science.gov (United States)

    Moore, Carleton J.

    1988-01-01

    The SRM propellant and polymer materials structural test program has potentially wide application to the testing and structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. The test program will provide a basis for characterization of the dynamic failure criteria for Solid Rocket Motor (SRM) propellant, insulation, inhibitor and liners. This experimental investigation will also endeavor to obtain a consistent complete set of materials test data. This test will be used to improve and revise the presently used theoretical math models for SRM propellant, insulators, inhibitor, liners, and O-ring seals.

  8. Modified computation of the nozzle damping coefficient in solid rocket motors

    Science.gov (United States)

    Liu, Peijin; Wang, Muxin; Yang, Wenjing; Gupta, Vikrant; Guan, Yu; Li, Larry K. B.

    2018-02-01

    In solid rocket motors, the bulk advection of acoustic energy out of the nozzle constitutes a significant source of damping and can thus influence the thermoacoustic stability of the system. In this paper, we propose and test a modified version of a historically accepted method of calculating the nozzle damping coefficient. Building on previous work, we separate the nozzle from the combustor, but compute the acoustic admittance at the nozzle entry using the linearized Euler equations (LEEs) rather than with short nozzle theory. We compute the combustor's acoustic modes also with the LEEs, taking the nozzle admittance as the boundary condition at the combustor exit while accounting for the mean flow field in the combustor using an analytical solution to Taylor-Culick flow. We then compute the nozzle damping coefficient via a balance of the unsteady energy flux through the nozzle. Compared with established methods, the proposed method offers competitive accuracy at reduced computational costs, helping to improve predictions of thermoacoustic instability in solid rocket motors.

  9. A Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

    Science.gov (United States)

    Yang, H. Q.; West, Jeff

    2014-01-01

    A capability to couple NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This paper summarizes the efforts in applying the installed coupling software to demonstrate/investigate fluid-structure interaction (FSI) between pressure wave and flexible inhibitor inside reusable solid rocket motor (RSRM). First a unified governing equation for both fluid and structure is presented, then an Eulerian-Lagrangian framework is described to satisfy the interfacial continuity requirements. The features of fluid solver, Loci/CHEM and structural solver, CoBi, are discussed before the coupling methodology of the solvers is described. The simulation uses production level CFD LES turbulence model with a grid resolution of 80 million cells. The flexible inhibitor is modeled with full 3D shell elements. Verifications against analytical solutions of structural model under steady uniform pressure condition and under dynamic condition of modal analysis show excellent agreements in terms of displacement distribution and eigen modal frequencies. The preliminary coupled result shows that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor.

  10. O-ring sealing verification for the space shuttle redesign solid rocket motor

    Science.gov (United States)

    Lach, Cynthia L.

    1989-01-01

    As a part of the redesign of the Space Shuttle Solid Rocket Motor, the field and nozzle-to-case joints were redesigned to minimize the dynamic flexure caused by internal motor pressurization during ignition. The O-ring seals and glands for the joints were designed to accommodate both structural deflections and to promote pressure assistance. A test program was conducted to determine if a fluorocarbon elastomeric O-ring could meet this criteria in the redesigned gland. Resiliency tests were used to investigate the O-ring response to gap motion while static seal tests were used to verify design criteria of pressure assistance for sealing. All tests were conducted in face seal fixtures mounted in servo-hydraulic test machines. The resiliency of the O-ring was found to be extremely sensitive to the effects of temperature. The External Tank/Solid Rocket Booster attach strut loads had a negligible affect on the ability of the O-ring to track the simulated SRB field joint deflection. In the static pressure-assisted seal tests, as long as physical contact was maintained between the O-ring and the gland sealing surface, pressure assistance induced instantaneous sealing.

  11. An Improved Approach for Hybrid Rocket Injection System Design

    OpenAIRE

    M. Invigorito; G. Elia; M. Panelli

    2016-01-01

    Hybrid propulsion combines beneficial properties of both solid and liquid rockets, such as multiple restarts, throttability as well as simplicity and reduced costs. A nitrous oxide (N2O)/paraffin-based hybrid rocket engine demonstrator is currently under development at the Italian Aerospace Research Center (CIRA) within the national research program HYPROB, funded by the Italian Ministry of Research. Nitrous oxide belongs to the class of self-pressurizing propellants that exhibit a high vapor...

  12. Simulation of reactive polydisperse sprays strongly coupled to unsteady flows in solid rocket motors: Efficient strategy using Eulerian Multi-Fluid methods

    Science.gov (United States)

    Sibra, A.; Dupays, J.; Murrone, A.; Laurent, F.; Massot, M.

    2017-06-01

    In this paper, we tackle the issue of the accurate simulation of evaporating and reactive polydisperse sprays strongly coupled to unsteady gaseous flows. In solid propulsion, aluminum particles are included in the propellant to improve the global performances but the distributed combustion of these droplets in the chamber is suspected to be a driving mechanism of hydrodynamic and acoustic instabilities. The faithful prediction of two-phase interactions is a determining step for future solid rocket motor optimization. When looking at saving computational ressources as required for industrial applications, performing reliable simulations of two-phase flow instabilities appears as a challenge for both modeling and scientific computing. The size polydispersity, which conditions the droplet dynamics, is a key parameter that has to be accounted for. For moderately dense sprays, a kinetic approach based on a statistical point of view is particularly appropriate. The spray is described by a number density function and its evolution follows a Williams-Boltzmann transport equation. To solve it, we use Eulerian Multi-Fluid methods, based on a continuous discretization of the size phase space into sections, which offer an accurate treatment of the polydispersion. The objective of this paper is threefold: first to derive a new Two Size Moment Multi-Fluid model that is able to tackle evaporating polydisperse sprays at low cost while accurately describing the main driving mechanisms, second to develop a dedicated evaporation scheme to treat simultaneously mass, moment and energy exchanges with the gas and between the sections. Finally, to design a time splitting operator strategy respecting both reactive two-phase flow physics and cost/accuracy ratio required for industrial computations. Using a research code, we provide 0D validations of the new scheme before assessing the splitting technique's ability on a reference two-phase flow acoustic case. Implemented in the industrial

  13. Coupled Solid Rocket Motor Ballistics and Trajectory Modeling for Higher Fidelity Launch Vehicle Design

    Science.gov (United States)

    Ables, Brett

    2014-01-01

    Multi-stage launch vehicles with solid rocket motors (SRMs) face design optimization challenges, especially when the mission scope changes frequently. Significant performance benefits can be realized if the solid rocket motors are optimized to the changing requirements. While SRMs represent a fixed performance at launch, rapid design iterations enable flexibility at design time, yielding significant performance gains. The streamlining and integration of SRM design and analysis can be achieved with improved analysis tools. While powerful and versatile, the Solid Performance Program (SPP) is not conducive to rapid design iteration. Performing a design iteration with SPP and a trajectory solver is a labor intensive process. To enable a better workflow, SPP, the Program to Optimize Simulated Trajectories (POST), and the interfaces between them have been improved and automated, and a graphical user interface (GUI) has been developed. The GUI enables real-time visual feedback of grain and nozzle design inputs, enforces parameter dependencies, removes redundancies, and simplifies manipulation of SPP and POST's numerous options. Automating the analysis also simplifies batch analyses and trade studies. Finally, the GUI provides post-processing, visualization, and comparison of results. Wrapping legacy high-fidelity analysis codes with modern software provides the improved interface necessary to enable rapid coupled SRM ballistics and vehicle trajectory analysis. Low cost trade studies demonstrate the sensitivities of flight performance metrics to propulsion characteristics. Incorporating high fidelity analysis from SPP into vehicle design reduces performance margins and improves reliability. By flying an SRM designed with the same assumptions as the rest of the vehicle, accurate comparisons can be made between competing architectures. In summary, this flexible workflow is a critical component to designing a versatile launch vehicle model that can accommodate a volatile

  14. Formulation and Testing of Paraffin-Based Solid Fuels Containing Energetic Additives for Hybrid Rockets

    Science.gov (United States)

    Larson, Daniel B.; Boyer, Eric; Wachs,Trevor; Kuo, Kenneth K.; Story, George

    2012-01-01

    Many approaches have been considered in an effort to improve the regression rate of solid fuels for hybrid rocket applications. One promising method is to use a fuel with a fast burning rate such as paraffin wax; however, additional performance increases to the fuel regression rate are necessary to make the fuel a viable candidate to replace current launch propulsion systems. The addition of energetic and/or nano-sized particles is one way to increase mass-burning rates of the solid fuels and increase the overall performance of the hybrid rocket motor.1,2 Several paraffin-based fuel grains with various energetic additives (e.g., lithium aluminum hydride (LiAlH4) have been cast in an attempt to improve regression rates. There are two major advantages to introducing LiAlH4 additive into the solid fuel matrix: 1) the increased characteristic velocity, 2) decreased dependency of Isp on oxidizer-to-fuel ratio. The testing and characterization of these solid-fuel grains have shown that continued work is necessary to eliminate unburned/unreacted fuel in downstream sections of the test apparatus.3 Changes to the fuel matrix include higher melting point wax and smaller energetic additive particles. The reduction in particle size through various methods can result in more homogeneous grain structure. The higher melting point wax can serve to reduce the melt-layer thickness, allowing the LiAlH4 particles to react closer to the burning surface, thus increasing the heat feedback rate and fuel regression rate. In addition to the formulation of LiAlH4 and paraffin wax solid-fuel grains, liquid additives of triethylaluminum and diisobutylaluminum hydride will be included in this study. Another promising fuel formulation consideration is to incorporate a small percentage of RDX as an additive to paraffin. A novel casting technique will be used by dissolving RDX in a solvent to crystallize the energetic additive. After dissolving the RDX in a solvent chosen for its compatibility

  15. Simulation of two-dimensional interior ballistics model of solid propellant electrothermal-chemical launch with discharge rod plasma generator

    Directory of Open Access Journals (Sweden)

    Yan-jie Ni

    2017-08-01

    Full Text Available Instead of the capillary plasma generator (CPG, a discharge rod plasma generator (DRPG is used in the 30 mm electrothermal-chemical (ETC gun to improve the ignition uniformity of the solid propellant. An axisymmetric two-dimensional interior ballistics model of the solid propellant ETC gun (2D-IB-SPETCG is presented to describe the process of the ETC launch. Both calculated pressure and projectile muzzle velocity accord well with the experimental results. The feasibility of the 2D-IB-SPETCG model is proved. Depending on the experimental data and initial parameters, detailed distribution of the ballistics parameters can be simulated. With the distribution of pressure and temperature of the gas phase and the propellant, the influence of plasma during the ignition process can be analyzed. Because of the radial flowing plasma, the propellant in the area of the DRPG is ignited within 0.01 ms, while all propellant in the chamber is ignited within 0.09 ms. The radial ignition delay time is much less than the axial delay time. During the ignition process, the radial pressure difference is less than 5  MPa at the place 0.025 m away from the breech. The radial ignition uniformity is proved. The temperature of the gas increases from several thousand K (conventional ignition to several ten thousand K (plasma ignition. Compare the distribution of the density and temperature of the gas, we know that low density and high temperature gas appears near the exits of the DRPG, while high density and low temperature gas appears at the wall near the breech. The simulation of the 2D-IB-SPETCG model is an effective way to investigate the interior ballistics process of the ETC launch. The 2D-IB-SPETC model can be used for prediction and improvement of experiments.

  16. Monitoring Microstructural Evolution and Crack Formation in a Solid Propellant under Incremental Strain Condition- Using Digital Radiography X-Ray Techniques

    National Research Council Canada - National Science Library

    Liu, C

    2004-01-01

    In this study, the change of microstructure and the formation of cracks in a solid propellant under an incremental strain loading condition were investigated using digital radiography x-ray techniques...

  17. 3-D Flash Lidar Performance in Flight Testing on the Morpheus Autonomous, Rocket-Propelled Lander to a Lunar-Like Hazard Field

    Science.gov (United States)

    Roback, Vincent E.; Amzajerdian, Farzin; Bulyshev, Alexander E.; Brewster, Paul F.; Barnes, Bruce W.

    2016-01-01

    For the first time, a 3-D imaging Flash Lidar instrument has been used in flight to scan a lunar-like hazard field, build a 3-D Digital Elevation Map (DEM), identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The flight tests served as the TRL 6 demo of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) system and included launch from NASA-Kennedy, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range. The ALHAT project developed a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The Flash Lidar is a second generation, compact, real-time, air-cooled instrument. Based upon extensive on-ground characterization at flight ranges, the Flash Lidar was shown to be capable of imaging hazards from a slant range of 1 km with an 8 cm range precision and a range accuracy better than 35 cm, both at 1-delta. The Flash Lidar identified landing hazards as small as 30 cm from the maximum slant range which Morpheus could achieve (450 m); however, under certain wind conditions it was susceptible to scintillation arising from air heated by the rocket engine and to pre-triggering on a dust cloud created during launch and transported down-range by wind.

  18. Hazard Studies for Solid Propellant Rocket Motors (Etude des Risque pour les Moteurs-Fusees a Propergols Solides)

    Science.gov (United States)

    1990-09-01

    stores. Those adjacent stores can be on the launch platform (tank, ship, anrft), in somne transportaion configuration (rail cars, trucks), in storage...the use of transparent tubes (Lexan), continuous access streak cameras, continuous access framing camera, and flash x-ray instrumentation were used

  19. Testing and environmental exposure of parachute materials for the solid rocket booster decelerator subsystem

    Science.gov (United States)

    Tannehill, B. K.

    1978-01-01

    Static tests and evaluation of nonmetallic materials for use in parachutes for recovery of solid rocket boosters used in the space shuttle program are reported. Literature survey and manufacturer and vendor contacts led to the choice of nylon as the fabric most capable of withstanding the extreme loads and environmental conditions during repeated use. The material tests included rupture strength, elongation, abrasion resistance, shrinkage, environmental exposure, and degradation levels. Rinsing and drying procedures were also investigated and a salt-free level for nylon recommended in preparation for reuse. In all possible cases, worst-case conditions were used (e.g., inflation loads, seawater exposure for 3 days per drop-recovery, etc.).

  20. Pressure oscillations and instability of working processes in the combustion chambers of solid rocket motors

    Science.gov (United States)

    Emelyanov, V. N.; Teterina, I. V.; Volkov, K. N.; Garkushev, A. U.

    2017-06-01

    Metal particles are widely used in space engineering to increase specific impulse and to supress acoustic instability of intra-champber processes. A numerical analysis of the internal injection-driven turbulent gas-particle flows is performed to improve the current understanding and modeling capabilities of the complex flow characteristics in the combustion chambers of solid rocket motors (SRMs) in presence of forced pressure oscillations. The two-phase flow is simulated with a combined Eulerian-Lagrangian approach. The Reynolds-averaged Navier-Stokes equations and transport equations of k - ε model are solved numerically for the gas. The particulate phase is simulated through a Lagrangian deterministic and stochastic tracking models to provide particle trajectories and particle concentration. The results obtained highlight the crucial significance of the particle dispersion in turbulent flowfield and high potential of statistical methods. Strong coupling between acoustic oscillations, vortical motion, turbulent fluctuations and particle dynamics is observed.

  1. Space Shuttle Solid Rocket Motor Plume Pressure and Heat Rate Measurements

    Science.gov (United States)

    vonEckroth, Wulf; Struchen, Leah; Trovillion, Tom; Perez, Ravael; Nereolich, Shaun; Parlier, Chris

    2012-01-01

    The Solid Rocket Booster (SRB) Main Flame Deflector (MFD) at Launch Complex 39A was instrumented with sensors to measure heat rates, pressures, and temperatures on the last three Space Shuttle launches. Because the SRB plume is hot and erosive, a robust Tungsten Piston Calorimeter was developed to compliment the measurements made by off-the-shelf sensors. Witness materials were installed and their melting and erosion response to the Mach 2 / 4500 F / 4-second duration plume was observed. The data show that the specification document used for the design of the MFD thermal protection system over-predicted heat rates by a factor of 3 and under-predicted pressures by a factor of 2. These findings will be used to baseline NASA Computational Fluid Dynamics models and develop innovative MFD designs for the Space Launch System (SLS) before this vehicle becomes operational in 2017.

  2. Optimization of the stand for test of hybrid rocket engines of solid fuel

    Directory of Open Access Journals (Sweden)

    Zolotorev Nikolay

    2017-01-01

    Full Text Available In the paper the laboratory experimental stand of the hybrid rocket engine of solid fuel to study ballistic parameters of the engine at burning of high-energy materials in flow of hot gas is presented. Mixture of air with nitrogen with a specified content of active oxygen is used as a gaseous oxidizer. The experimental stand has modular design and consists of system of gas supply, system of heating of gas, system for monitoring gas parameters, to which a load cell with a model engine was connected. The modular design of the stand allows to change its configuration under specific objective. This experimental stand allows to conduct a wide range of the pilot studies at interaction of a hot stream of gas with samples high-energy materials.

  3. Analytical flow/thermal modeling of combustion gas flows in Redesigned Solid Rocket Motor test joints

    Science.gov (United States)

    Woods, G. H.; Knox, E. C.; Pond, J. E.; Bacchus, D. L.; Hengel, J. E.

    1992-07-01

    A one-dimensional analytical tool, TOPAZ (Transient One-dimensional Pipe flow AnalyZer), was used to model the flow characteristics of hot combustion gases through Redesigned Solid Rocket Motor (RSRM) joints and to compute the resultant material surface temperatures and o-ring seal erosion of the joints. The capabilities of the analytical tool were validated with test data during the Seventy Pound Charge (SPC) motor test program. The predicted RSRM joint thermal response to ignition transients was compared with test data for full-scale motor tests. The one-dimensional analyzer is found to be an effective tool for simulating combustion gas flows in RSRM joints and for predicting flow and thermal properties.

  4. Workers in the VAB test SRB cables on STS-98 solid rocket boosters

    Science.gov (United States)

    2001-01-01

    KENNEDY SPACE CENTER, Fla. -- United Space Alliance SRB technician Richard Bruns attaches a cable end cover to a cable pulled from the solid rocket booster on Space Shuttle Atlantis. The Shuttle was rolled back from Launch Pad 39A in order to conduct tests on the SRB cables. A prior extensive evaluation of NASA'''s SRB cable inventory on the shelf revealed conductor damage in four (of about 200) cables. Shuttle managers decided to prove the integrity of the system tunnel cables already on Atlantis before launching. Workers are conducting inspections, making continuity checks and conducting X-ray analysis on the cables. The launch has been rescheduled no earlier than Feb. 6.

  5. SRB/SLEEC (Solid Rocket Booster/Shingle Lap Extendible Exit Cone) feasibility study, volume 1

    Science.gov (United States)

    Baker, William H., Jr.

    1986-01-01

    A preliminary design and analysis was completed for a SLEEC (Shingle Lap Extendible Exit Cone) which could be incorporated on the Space Transportation System (STS) Solid Rocket Booster (SRB). Studies were completed which predicted weights and performance increases and development plans were prepared for the full-scale bench and static test of SLEEC. In conjunction with the design studies, a series of supporting analyses were performed to assure the validity and feasibility of performance, fabrication, cost, and reliability for the selected design. The feasibility and required amounts of bench, static firing, and flight tests considered necessary for the successful incorporation of SLEEC on the Shuttle SRBs were determined. Preliminary plans were completed which define both a follow on study effort and a development program.

  6. NDE of thermal protection system for space shuttle solid rocket booster

    Science.gov (United States)

    Myers, R. S.

    1990-01-01

    Potential nondestructive test (NDE) methods were evaluated for detecting debonds and weak bonds in the thermal protection system (TPS) for the space shuttle solid rocket boosters. The primary thermal protection material is a sprayable, thick epoxy coating that is filled with lightweight and thermal insulating materials. Test panels were fabricated with a wide variety of hidden realistic defects, including contact debonds and weak bonds. Nondestructive test results were obtained. Candidate NDE methods evaluated for booster production applications include laser interferometry (e.g., electronic shearography), infrared thermography, radiography (e.g., computed tomography), acousto-ultrasonics, mechanical/acoustic impedance, ultrasonics, acoustic emission, and the tap test. Capabilities, advantages, disadvantages, and relative performances in defect detection of each test method for TPS bonding applications are reported. Electronic shearography NDE was technically the superior method for detecting debonds.

  7. Flight Test Results of Rocket-Propelled Buffet-Research Models Having 45 Degree Sweptback Wings and 45 Degree Sweptback Tails Located in the Wing Chord Plane

    Science.gov (United States)

    Mason, Homer P.

    1953-01-01

    Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.

  8. Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

    Science.gov (United States)

    Bellman, Donald R; Humphrey, Jack C

    1948-01-01

    Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

  9. Oxidation/Biodegradation of Solid Propellants From a 4.2-Inch Chemical Round

    Science.gov (United States)

    2004-11-16

    Microtox (MTX) assay. The MICROTOX assay exposes a bioluminescent marine bacterium ( Vibrio fischeri ) to a sample of unknown toxicity and measuring the...the Microtox values indicates a decrease in the toxicity of the propellant feed to the Vibrio fischeri . This data can be used, as an indicator as to...not only non-toxic but has become nutritive to Vibrio fischeri . The overall effectiveness of the strategy 1 approach for both propellant types

  10. Ballistic properties and burning behaviour of an ammonium perchlorate/guanidine nitrate/sodium nitrate airbag solid propellant

    Energy Technology Data Exchange (ETDEWEB)

    A. Ulas; G.A. Risha; K.K. Kuo [Middle East Technical University, Ankara (Turkey). Department of Mechanical Engineering

    2006-10-15

    An experimental investigation on the determination of ballistic properties and burning behavior of a composite solid propellant for airbag application was conducted. The experimental results were obtained using a high-pressure optical strand burner. Steady-state burning rates were determined for a pressure range of 20.8-41.5 MPa and initial propellant temperatures of -30 to +80{sup o}C. For the pressure and temperature ranges tested, the temperature sensitivity was on the order of 1 x 10{sup -3} K{sup -1}. The pressure exponent was found to be a function of the initial propellant temperature and was 0.75 at 25{sup o}C. The activation energy and the pre-exponential factor of the Arrhenius equation are 2.735 kcal/mol and 15.06 cm/s, respectively. The pressure deflagration limit for this propellant was found to be in the range of 8.37-8.72 MPa. During combustion, small condensed-phase spherical particles were ejected from the burning surface. The size of the particles decreased with either increasing the pressure or the initial propellant temperature. For pressures below 41.4 MPa, average particle size was on the order of 900 {mu}m, and at 84.4 MPa, the bead size was much smaller, on the order of 300 {mu}m. A chemical analysis on these particles using both the ESEM and the X-ray diffraction method indicated that the material of the beads was mostly sodium chloride with a small amount of silicon-containing compounds. 6 refs., 17 figs.

  11. Hybrid rocket motor testing at Nammo Raufoss A/S

    Science.gov (United States)

    Rønningen, Jan-Erik; Kubberud, Nils

    2005-08-01

    Hybrid rocket motor technology and the use of hybrid rockets have gained increased interest in recent years in many countries. A typical hybrid rocket consists of a tank containing the oxidizer in either liquid or gaseous state connected to the combustion chamber containing an injector, inert solid fuel grain and nozzle. Nammo Raufoss A/S has for almost 40 years designed and produced high-performance solid propellant rocket motors for many military missile systems as well as solid propellant rocket motors for civil space use. In 2003 an in-house technology program was initiated to investigate and study hybrid rocket technology. On 23 September 2004 the first in-house designed hybrid test rocket motor was static test fired at Nammo Raufoss Test Center. The oxidizer was gaseous oxygen contained in a tank pressurized to 10MPa, flow controlled through a sonic orifice into the combustion chamber containing a multi port radial injector and six bore cartridge-loaded fuel grain containing a modified HTPB fuel composition. The motor was ignited using a non-explosive heated wire. This paper will present what has been achieved at Nammo Raufoss since the start of the program.

  12. Arc Jet Test and Analysis of Asbestos Free Solid Rocket Motor Nozzle Dome Ablative Materials

    Science.gov (United States)

    Clayton, J. Louie

    2017-01-01

    Asbestos free solid motor internal insulation samples were recently tested at the MSFC Hyperthermal Arc Jet Facility. Objectives of the test were to gather data for solid rocket motor analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Active instrumentation included use of total calorimeters, in-depth thermocouples, and a surface pyrometer for in-situ surface temperature measurement. Post-test sample forensics involved determination of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

  13. The Thermal State Computational Research of the Low-Thrust Oxygen-Methane Gaseous-Propellant Rocket Engine in the Pulse Mode of Operation

    Directory of Open Access Journals (Sweden)

    O. A. Vorozheeva

    2014-01-01

    Full Text Available Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is therefore necessary in the development phase of a new product. Accordingly, the aim of this work is the mathematical modeling and computational study of the thermal state of gaseous oxygen-methane propellant RDMT operating in pulse mode.In this paper we consider a model RDTM working on gaseous propellants oxygen-methane in pulse mode.To calculate the temperature field of the chamber wall of model RDMT under consideration is used the mathematical model of non-stationary heat conduction in a two-dimensional axisymmetric formulation that takes into account both the axial heat leakages and the nonstationary processes occurring inside the chamber during pulse operation of RDMT.As a result of numerical study of the thermal state of model RDMT, are obtained the temperature fields during engine operation based on convective, conductive, and radiative mechanisms of heat transfer from the combustion products to the wall.It is shown that the elements of flanges of combustion chamber of model RDMT act as heat sinks structural elements. Temperatures in the wall of the combustion chamber during the engine mode of operation are considered relatively low.Raised temperatures can also occur in the mixing head in the feeding area of the oxidant into the combustion chamber.During engine operation in the area forming the critical section, there is an intensive heating of a wall, which can result in its melting, which in turn will increase the minimum nozzle throat area and hence

  14. Mars Integrated Propellant Production System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Integrated Mars In-Situ Propellant Production System (IMISPPS) is an end-to-end system that will produce rocket propellant on Mars from CO2 in the Martian...

  15. U.S. Army Armament Research, Development and Engineering Center Grain Evaluation Software to Numerically Predict Linear Burn Regression for Solid Propellant Grain Geometries

    Science.gov (United States)

    2017-10-01

    ENGINEERING CENTER GRAIN EVALUATION SOFTWARE TO NUMERICALLY PREDICT LINEAR BURN REGRESSION FOR SOLID PROPELLANT GRAIN GEOMETRIES Brian...distribution is unlimited. AD U.S. ARMY ARMAMENT RESEARCH, DEVELOPMENT AND ENGINEERING CENTER Munitions Engineering Technology Center Picatinny...U.S. ARMY ARMAMENT RESEARCH, DEVELOPMENT AND ENGINEERING CENTER GRAIN EVALUATION SOFTWARE TO NUMERICALLY PREDICT LINEAR BURN REGRESSION FOR SOLID

  16. Solid Rocket Booster (SRB) - Evolution and Lessons Learned During the Shuttle Program

    Science.gov (United States)

    Kanner, Howard S.; Freeland, Donna M.; Olson, Derek T.; Wood, T. David; Vaccaro, Mark V.

    2011-01-01

    The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Obsolescence issues occasionally required component recertification. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. Assembly and integration of the booster subsystems was a unique process and acceptance testing of reused hardware components was required for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed. The postflight assessment process was quite detailed and a

  17. Real-Time Measurements of Aft Dome Insulation Erosion on Space Shuttle Reusable Solid Rocket Motor

    Science.gov (United States)

    McWhorter, Bruce; Ewing, Mark; Albrechtsen, Kevin; Noble, Todd; Longaker, Matt

    2004-01-01

    Real-time erosion of aft dome internal insulation was measured with internal instrumentation on a static test of a lengthened version of the Space Shuffle Reusable Solid Rocket Motor (RSRM). This effort marks the first time that real-time aft dome insulation erosion (Le., erosion due to the combined effects of thermochemical ablation and mechanical abrasion) was measured in this kind of large motor static test [designated as Engineering Test Motor number 3 (ETM3)I. This paper presents data plots of the erosion depth versus time. The data indicates general erosion versus time behavior that is in contrast to what would be expected from earlier analyses. Engineers have long known that the thermal environment in the aft dome is severe and that the resulting aft dome insulation erosion is significant. Models of aft dome erosion involve a two-step process of computational fluid dynamics (CFD) modeling and material ablation modeling. This modeling effort is complex. The time- dependent effects are difficult to verify with only prefire and postfire insulation measurements. Nozzle vectoring, slag accumulation, and changing boundary conditions will affect the time dependence of aft dome erosion. Further study of this data and continued measurements on future motors will increase our understanding of the aft dome flow and erosion environment.

  18. Development of an On-board Failure Diagnostics and Prognostics System for Solid Rocket Booster

    Science.gov (United States)

    Smelyanskiy, Vadim N.; Luchinsky, Dmitry G.; Osipov, Vyatcheslav V.; Timucin, Dogan A.; Uckun, Serdar

    2009-01-01

    We develop a case breach model for the on-board fault diagnostics and prognostics system for subscale solid-rocket boosters (SRBs). The model development was motivated by recent ground firing tests, in which a deviation of measured time-traces from the predicted time-series was observed. A modified model takes into account the nozzle ablation, including the effect of roughness of the nozzle surface, the geometry of the fault, and erosion and burning of the walls of the hole in the metal case. The derived low-dimensional performance model (LDPM) of the fault can reproduce the observed time-series data very well. To verify the performance of the LDPM we build a FLUENT model of the case breach fault and demonstrate a good agreement between theoretical predictions based on the analytical solution of the model equations and the results of the FLUENT simulations. We then incorporate the derived LDPM into an inferential Bayesian framework and verify performance of the Bayesian algorithm for the diagnostics and prognostics of the case breach fault. It is shown that the obtained LDPM allows one to track parameters of the SRB during the flight in real time, to diagnose case breach fault, and to predict its values in the future. The application of the method to fault diagnostics and prognostics (FD&P) of other SRB faults modes is discussed.

  19. Flow Simulation of Solid Rocket Motors. 2; Sub-Scale Air Flow Simulation of Port Flows

    Science.gov (United States)

    Yeh, Y. P.; Ramandran, N.; Smith, A. W.; Heaman, J. P.

    2000-01-01

    The injection-flow issuing from a porous medium in the cold-flow simulation of internal port flows in solid rocket motors is characterized by a spatial instability termed pseudoturbulence that produces a rather non-uniform (lumpy) injection-velocity profile. The objective of this study is to investigate the interaction between the injection- and the developing axial-flows. The findings show that this interaction generally weakens the lumpy injection profile and affects the subsequent development of the axial flow. The injection profile is found to depend on the material characteristics, and the ensuing pseudoturbulence is a function of the injection velocity, the axial position and the distance from the porous wall. The flow transition (from laminar to turbulent) of the axial-flow is accelerated in flows emerging from smaller pores primarily due to the higher pseudoturbulence produced by the smaller pores in comparison to that associated with larger pores. In flows with rather uniform injection-flow profiles (weak or no pseudoturbulence), the axial and transverse velocity components in the porous duct are found to satisfy the sine/cosine analytical solutions derived from inviscid assumptions. The transition results from the present study are compared with previous results from surveyed literature, and detailed flow development measurements are presented in terms of the blowing fraction, and characterizing Reynolds numbers.

  20. Flow Simulation of Solid Rocket Motors. 1; Injection Induced Water-Flow Tests from Porous Media

    Science.gov (United States)

    Ramachandran, N.; Yeh, Y. P.; Smith, A. W.; Heaman, J. P.

    1999-01-01

    Prior to selecting a proper porous material for use in simulating the internal port flow of a solid rocket motor (SRM), in cold-flow testing, the flow emerging from porous materials is experimentally investigated. The injection-flow emerging from a porous matrix always exhibits a lumpy velocity profile that is spatially stable and affects the development of the longitudinal port flow. This flow instability, termed pseudoturbulence, is an inherent signature of the porous matrix and is found to generally increase with the wall porosity and with the injection flow rate. Visualization studies further show that the flow from porous walls made from shaving-type material (sintered stainless-steel) exhibits strong recirculation zones that are conspicuously absent in walls made from nodular or spherical material (sintered bronze). Detailed flow visualization observations and hot-film measurements are reported from tests of injection-flow and a coupled cross-flow from different porous wall materials. Based on the experimental data, discussion is provided on the choice of suitable material for SRM model testing while addressing the consequences and shortcomings from such a test.

  1. Laser Shearography Inspection of TPS (Thermal Protection System) Cork on RSRM (Reusable Solid Rocket Motors)

    Science.gov (United States)

    Lingbloom, Mike; Plaia, Jim; Newman, John

    2006-01-01

    Laser Shearography is a viable inspection method for detection of de-bonds and voids within the external TPS (thermal protection system) on to the Space Shuttle RSRM (reusable solid rocket motors). Cork samples with thicknesses up to 1 inch were tested at the LTI (Laser Technology Incorporated) laboratory using vacuum-applied stress in a vacuum chamber. The testing proved that the technology could detect cork to steel un-bonds using vacuum stress techniques in the laboratory environment. The next logical step was to inspect the TPS on a RSRM. Although detailed post flight inspection has confirmed that ATK Thiokol's cork bonding technique provides a reliable cork to case bond, due to the Space Shuttle Columbia incident there is a great interest in verifying bond-lines on the external TPS. This interest provided and opportunity to inspect a RSRM motor with Laser Shearography. This paper will describe the laboratory testing and RSRM testing that has been performed to date. Descriptions of the test equipment setup and techniques for data collection and detailed results will be given. The data from the test show that Laser Shearography is an effective technology and readily adaptable to inspect a RSRM.

  2. Oxidation/Biodegradation of Solid Propellants Used in Legacy Chemical Rounds

    Science.gov (United States)

    2007-08-01

    metal parts clean out. Complete characterization of the compounds in the propellant hydrolysate is difficult due to the sheer number of products found...bacterial identification. Biomass suspension was extracted from the M28 process I and 2 reactors, and plated raw material. All incubation steps occurred...20,616 Dimethyl Phenanthroline 1000 7,657 Feed 1834 Phenyl-Benzotriazole 618 Decanal 1075 M28-1 3-hr 1,4-Dioxane 25 Furfural 26 treated Heptanal 18 43

  3. Lidar Sensor Performance in Closed-Loop Flight Testing of the Morpheus Rocket-Propelled Lander to a Lunar-Like Hazard Field

    Science.gov (United States)

    Roback, V. Eric; Pierrottet, Diego F.; Amzajerdian, Farzin; Barnes, Bruce W.; Bulyshev, Alexander E.; Hines, Glenn D.; Petway, Larry B.; Brewster, Paul F.; Kempton, Kevin S.

    2015-01-01

    For the first time, a suite of three lidar sensors have been used in flight to scan a lunar-like hazard field, identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The lidar sensors and GN&C system are part of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) project which has been seeking to develop a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The 3-D imaging Flash Lidar is a second generation, compact, real-time, aircooled instrument developed from a number of components from industry and NASA and is used as part of the ALHAT Hazard Detection System (HDS) to scan the hazard field and build a 3-D Digital Elevation Map (DEM) in near-real time for identifying safe sites. The Flash Lidar is capable of identifying a 30 cm hazard from a slant range of 1 km with its 8 cm range precision (1-s). The Flash Lidar is also used in Hazard Relative Navigation (HRN) to provide position updates down to a 250m slant range to the ALHAT navigation filter as it guides Morpheus to the safe site. The Navigation Doppler Lidar (NDL) system has been developed within NASA to provide velocity measurements with an accuracy of 0.2 cm/sec and range measurements with an accuracy of 17 cm both from a maximum range of 2,200 m to a minimum range of several meters above the ground. The NDLâ€"TM"s measurements are fed into the ALHAT navigation filter to provide lander guidance to the safe site. The Laser Altimeter (LA), also developed within NASA, provides range measurements with an accuracy of 5 cm from a maximum operational range of 30 km down to 1 m and, being a separate sensor from the Flash Lidar, can provide range along a separate vector. The LA measurements are also fed

  4. Solid Rocket Motor for Ultralow Temperature Operation During the Mars Sample Return Mission Project

    Data.gov (United States)

    National Aeronautics and Space Administration — A small Mars (or other celestial body) ascent vehicle is unlikely to achieve the necessary propellant fraction required to achieve orbit. Scaling down of liquid...

  5. The Limitations of Upper-Atmosphere Research Vehicles Powered by Current British Solid-Fuel Rockets

    National Research Council Canada - National Science Library

    King-Hele, D. G

    1953-01-01

    .... If current rockets are to be of value in upper-atmosphere research, therefore, they must either be used in a single-stage arrangement, or be launched well above sea level, perhaps from a balloon or mountain top...

  6. Pulsed plasma solid propellant microthruster for the synchronous meteorological satellite. Task 4: Engineering model fabrication and test report

    Science.gov (United States)

    Guman, W. J. (Editor)

    1972-01-01

    Two flight prototype solid propellant pulsed plasma microthruster propulsion systems for the SMS satellite were fabricated, assembled and tested. The propulsion system is a completely self contained system requiring only three electrical inputs to operate: a 29.4 volt power source, a 28 volt enable signal and a 50 millsec long command fire signal that can be applied at any rate from 50 ppm to 110 ppm. The thrust level can be varied over a range 2.2 to 1 at constant impulse bit amplitude. By controlling the duration of the 28 volt enable either steady state thrust or a series of discrete impulse bits can be generated. A new technique of capacitor charging was implemented to reduce high voltage stress on energy storage capacitors.

  7. Solid Propellant Ignition and Other Unsteady Combustion Phenomena Induced by Radiation

    Science.gov (United States)

    1976-11-01

    transparent cell ; and limitations (radiation penetration and cold, chemically-inert atmosphere are not typical of rocket motor ignition). (2) the report...priori on simple physical grounds. To extract less immediate results is a cumbersome task and, moreover, is strongly limited by the requirement of...illumination study case. Fig. B2.b Integral balance check for the study case. .j , -H 44 E EEE 41:4 .p00 A A I Wr-4 N, I ~4.4 0 zz ’-4 U0 rT4 1-41 S

  8. The effect of solid phase reactions on the ballistic properties of propellants

    Science.gov (United States)

    Schmidt, W. G.

    1970-01-01

    The combustion of NH4ClO4 composite propellants has been studied between 15 and 3000 psi. The emphasis in the program has been on determining the mechanisms by which the fuel components influence the burning rate of the composites. In order to have flexibility in the choice and concentration of the fuel component all combustion experiments were performed with pressed power strands. The fuels studied included those which affected the combustion mechanism of the composite primarily through their effect on: (1) the oxidizer decomposition mechanism and (2) the composite surface temperature. The combustion of pure and doped NH4ClO4 was studied using both pressed powder strands and pressed end burning motor grains. The experimental approach has been essentially a chemical one with emphasis on perturbing those reactions which occur on or immediately adjacent to the surface (zone of influence) of the composite.

  9. Space shuttle solid rocket booster recovery system definition. Volume 2: SRB water impact Monte Carlo computer program, user's manual

    Science.gov (United States)

    1973-01-01

    The HD 220 program was created as part of the space shuttle solid rocket booster recovery system definition. The model was generated to investigate the damage to SRB components under water impact loads. The random nature of environmental parameters, such as ocean waves and wind conditions, necessitates estimation of the relative frequency of occurrence for these parameters. The nondeterministic nature of component strengths also lends itself to probabilistic simulation. The Monte Carlo technique allows the simultaneous perturbation of multiple independent parameters and provides outputs describing the probability distribution functions of the dependent parameters. This allows the user to determine the required statistics for each output parameter.

  10. Preflight transient dynamic analyses of B-52 aircraft carrying Space Shuttle solid rocket booster drop-test vehicle

    Science.gov (United States)

    Ko, W. L.; Schuster, L. S.

    1984-01-01

    This paper concerns the transient dynamic analysis of the B-52 aircraft carrying the Space Shuttle solid rocket booster drop test vehicle (SRB/DTV). The NASA structural analysis (NASTRAN) finite element computer program was used in the analysis. The B-52 operating conditions considered for analysis were (1) landing and (2) braking on aborted takeoff runs. The transient loads for the B-52 pylon front and rear hooks were calculated. The results can be used to establish the safe maneuver envelopes for the B-52 carrying the SRB/DTV in landings and brakings.

  11. Pre-flight transient dynamic analysis of B-52 carrying Space Shuttle solid rocket booster drop-test vehicle

    Science.gov (United States)

    Ko, W. L.; Schuster, L. S.

    1983-01-01

    This paper concerns the transient dynamic analysis of the B-52 aircraft carrying the Space Shuttle solid-rocket booster drop-test vehicle (SRB/DTV). The NASA structural analysis (NASTRAN) finite-element computer program was used in the analysis. The B-52 operating conditions considered for analysis were (1) landing and (2) braking on aborted takeoff runs. The transient loads for the B-52 pylon front and rear hooks were calculated. The results can be used to establish the safe maneuver envelopes for the B-52 carrying the SRB/DTV in landings and brakings.

  12. Direct electrical arc ignition of hybrid rocket motors

    Science.gov (United States)

    Judson, Michael I., Jr.

    Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

  13. Karl Poggensee - A widely unknown German rocket pioneer - The early years 1930-1934 - A chronology

    Science.gov (United States)

    Rohrwild, Karlheinz

    2017-09-01

    The rediscovered estate of Karl Poggensee allows to reproduce chronologically his rocket tests of the period 1930-1934 almost completely for the first time. Thrilled by the movie ;The Woman in the Moon; for the idea of space travel, he started as a student of Hinderburg-Polytechnikum (IAO), Oldenburg, to build his first solid-fuel rocket, producing his own propellant charges. Being a coming electrical engineer his main goal was not set up new record heights, but to provide his rockets with automatic measuring instruments, camera and parachute release systems. The optimization of this sequence was his main focus.

  14. Effect of temperature and gap opening rate on the resiliency of candidate solid rocket booster O-ring materials

    Science.gov (United States)

    Lach, Cynthia L.

    1992-01-01

    In the redesign of the Space Shuttle solid rocket motor following the Challenger accident, the field and nozzle-to-case joints were designed to minimize gap opening caused by internal motor pressurization during ignition. The O-ring seals and glands for these joints were designed both to accommodate structural deflections and to promote pressure assisted sealing. The resiliency behavior of several candidate O-ring materials was evaluated for the effects of temperature and gap opening rates. The performance of three of the elastomeric materials was tested under the specific redesign gap opening requirement. Dynamic flexure conditions unique to launch produce low frequency vibrations in the gap opening. The effect of these vibrations on the ability of the O-ring to maintain contact with the sealing surface was addressed. The resiliency of the O-ring materials was found to be extremely sensitive to variations in temperature and gap opening rate. The top three elastomeric materials tracked the simulated solid rocket booster (SRB) field joint deflection at 75 and 120 F. The external tank/SRB attach strut load vibrations had a negligible effect on the ability of the O-ring to track the simulated SRB field joint deflection.

  15. Propellant Readiness Level: A Methodological Approach to Propellant Characterization

    Science.gov (United States)

    Bossard, John A.; Rhys, Noah O.

    2010-01-01

    A methodological approach to defining propellant characterization is presented. The method is based on the well-established Technology Readiness Level nomenclature. This approach establishes the Propellant Readiness Level as a metric for ascertaining the readiness of a propellant or a propellant combination by evaluating the following set of propellant characteristics: thermodynamic data, toxicity, applications, combustion data, heat transfer data, material compatibility, analytical prediction modeling, injector/chamber geometry, pressurization, ignition, combustion stability, system storability, qualification testing, and flight capability. The methodology is meant to be applicable to all propellants or propellant combinations; liquid, solid, and gaseous propellants as well as monopropellants and propellant combinations are equally served. The functionality of the proposed approach is tested through the evaluation and comparison of an example set of hydrocarbon fuels.

  16. Innovative Swirl Injector for LOX and Hydrocarbon Propellants Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Gases trapped in the propellant feed lines of space-based rocket engines due to cryogenic propellant boil-off or pressurant ingestion can result in poor combustion...

  17. Large Payload Nuclear Rockets

    Science.gov (United States)

    1964-06-01

    CORE TIMP . 5013R SOLIDO FRACrON ,0.7 REFL MAT: No 25O00 EFL TEMP. • 4000R Tg a 3.0 CM sot a 1.0 w• 20000 -.. 0U.. 1,iI.000. 0 •0 4~100 \\ 080I0...Boiling and Super - critical Pressure States," ARS 1710-61, presented at American Rocket Society Propellants, Combustion, and Liquid Rockets Conference

  18. Metallic hydrogen: The most powerful rocket fuel yet to exist

    Energy Technology Data Exchange (ETDEWEB)

    Silvera, Isaac F [Lyman Laboratory of Physics, Harvard University, Cambridge MA 02138 (United States); Cole, John W, E-mail: silvera@physics.harvard.ed [NASA MSFC, Huntsville, AL 35801 (United States)

    2010-03-01

    Wigner and Huntington first predicted that pressures of order 25 GPa were required for the transition of solid molecular hydrogen to the atomic metallic phase. Later it was predicted that metallic hydrogen might be a metastable material so that it remains metallic when pressure is released. Experimental pressures achieved on hydrogen have been more than an order of magnitude higher than the predicted transition pressure and yet it remains an insulator. We discuss the applications of metastable metallic hydrogen to rocketry. Metastable metallic hydrogen would be a very light-weight, low volume, powerful rocket propellant. One of the characteristics of a propellant is its specific impulse, I{sub sp}. Liquid (molecular) hydrogen-oxygen used in modern rockets has an Isp of {approx}460s; metallic hydrogen has a theoretical I{sub sp} of 1700s. Detailed analysis shows that such a fuel would allow single-stage rockets to enter into orbit or carry economical payloads to the moon. If pure metallic hydrogen is used as a propellant, the reaction chamber temperature is calculated to be greater than 6000 K, too high for currently known rocket engine materials. By diluting metallic hydrogen with liquid hydrogen or water, the reaction temperature can be reduced, yet there is still a significant performance improvement for the diluted mixture.

  19. Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

    Directory of Open Access Journals (Sweden)

    Kalomba Mboyi

    2015-04-01

    Full Text Available A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the emitting material with the radioisotope decay heat and by powering the different valves of the plasma rocket engine with the same radioisotope decay heat using a radioisotope thermoelectric generator. This rocket engine is then benchmarked against a 1 N hydrazine thruster configuration operated on one of the Pleiades-HR-1 constellation spacecraft. A maximal specific impulse and power saving of respectively 529 s and 32% are achieved with helium as propellant. Its advantages are its power saving capability, high specific impulses and simultaneous ease of storage and restart. It can however be extremely voluminous and potentially hazardous. The Kabila rocket is found to bring great benefits to the existing spacecraft and further research should optimize its geometric characteristics and investigate the physical principals of its operation.

  20. Solid amine-boranes as high performance hypergolic hybrid rocket fuels

    Science.gov (United States)

    Pfeil, Mark A.

    Hypergolic hybrid rockets have the potential of providing systems that are simple, reliable, have high performance, and allow for energy management. Such a propulsion system can be applied to fields that need a single tactical motor with flexible mission requirements of either high speed to target or extended loitering. They also provide the possibility for alternative fast response dynamic altitude control systems if ignition delays are sufficiently short. Amines are the traditional fuel of choice when selecting a hypergolic combination as these tend to react readily with both nitric acid and dinitrogen tertroxide based oxidizers. It has been found that the addition of a borane adduct to an amine fuel tends to reduce the ignition delay by up to an order of magnitude with white fuming nitric acid (WFNA). The borane addition has resulted in fuels with very short ignition delays between 2-10 ms - the fastest times for an amine based fuel reacting with nitric acid based oxidizers. The incorporation of these amine-boranes, specifically ethylenediamine bisborane (EDBB), into various fuel binders has also been found to result in ignition delays between 3-10 ms - the fastest times again for amine based fuels. It was found that the addition of a borane to an amine increased theoretical performance of the amine resulting in high performance fuels. The amine-borane/fuel binder combinations also produced higher theoretical performance values than previously used hypergolic hybrid rockets. Some of the theoretical values are on par or higher than the current toxic liquid hypergolic fuels, making amine boranes an attractive replacement. The higher performing amine-borane/fuel binder combinations also have higher performance values than the traditional rocket fuels, excluding liquid hydrogen. Thus, amine-borane based fuels have the potential to influence various area in the rocket field. An EDBB/ferrocene/epoxy fuel was tested in a hypergolic hybrid with pure nitric acid as the

  1. Influence of Structural Parameters on the Performance of Vortex Valve Variable-Thrust Solid Rocket Motor

    Science.gov (United States)

    Wei, Xianggeng; Li, Jiang; He, Guoqiang

    2017-04-01

    The vortex valve solid variable thrust motor is a new solid motor which can achieve Vehicle system trajectory optimization and motor energy management. Numerical calculation was performed to investigate the influence of vortex chamber diameter, vortex chamber shape, and vortex chamber height of the vortex valve solid variable thrust motor on modulation performance. The test results verified that the calculation results are consistent with laboratory results with a maximum error of 9.5%. The research drew the following major conclusions: the optimal modulation performance was achieved in a cylindrical vortex chamber, increasing the vortex chamber diameter improved the modulation performance of the vortex valve solid variable thrust motor, optimal modulation performance could be achieved when the height of the vortex chamber is half of the vortex chamber outlet diameter, and the hot gas control flow could result in an enhancement of modulation performance. The results can provide the basis for establishing the design method of the vortex valve solid variable thrust motor.

  2. Testing of Environmentally Preferable Aluminum Pretreatments and Coating Systems for Use on Space Shuttle Solid Rocket Boosters (SRB)

    Science.gov (United States)

    Clayton, C.; Raley, R.; Zook, L.

    2001-01-01

    The solid rocket booster (SRB) has historically used a chromate conversion coating prior to protective finish application. After conversion coating, an organic paint system consisting of a chromated epoxy primer and polyurethane topcoat is applied. An overall systems approach was selected to reduce waste generation from the coatings application and removal processes. While the most obvious waste reduction opportunity involved elimination of the chromate conversion coating, several other coating system configurations were explored in an attempt to reduce the total waste. This paper will briefly discuss the use of a systems view to reduce waste generation from the coating process and present the results of the qualification testing of nonchromated aluminum pretreatments and alternate coating systems configurations.

  3. Challenges in Propellants and Combustion: 100 Years after Nobel

    National Research Council Canada - National Science Library

    Kuo, Kenneth

    1997-01-01

    .... The Topics covered include: chemical kinetics of propellant combustion, environmental considerations in combustion of solid and liquid propellants, commercial application in the combustion of energetic materials, effective...

  4. Advanced Flow Analysis Tools for Transient Solid Rocket Motor Simulations Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The challenges of designing, developing, and fielding man-rated propulsion systems continue to increase as NASA's mission moves forward with evolving solid...

  5. ELM - A SIMPLE TOOL FOR THERMAL-HYDRAULIC ANALYSIS OF SOLID-CORE NUCLEAR ROCKET FUEL ELEMENTS

    Science.gov (United States)

    Walton, J. T.

    1994-01-01

    ELM is a simple computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in nuclear thermal rockets. Written for the nuclear propulsion project of the Space Exploration Initiative, ELM evaluates the various heat transfer coefficient and friction factor correlations available for turbulent pipe flow with heat addition. In the past, these correlations were found in different reactor analysis codes, but now comparisons are possible within one program. The logic of ELM is based on the one-dimensional conservation of energy in combination with Newton's Law of Cooling to determine the bulk flow temperature and the wall temperature across a control volume. Since the control volume is an incremental length of tube, the corresponding pressure drop is determined by application of the Law of Conservation of Momentum. The size, speed, and accuracy of ELM make it a simple tool for use in fuel element parametric studies. ELM is a machine independent program written in FORTRAN 77. It has been successfully compiled on an IBM PC compatible running MS-DOS using Lahey FORTRAN 77, a DEC VAX series computer running VMS, and a Sun4 series computer running SunOS UNIX. ELM requires 565K of RAM under SunOS 4.1, 360K of RAM under VMS 5.4, and 406K of RAM under MS-DOS. Because this program is machine independent, no executable is provided on the distribution media. The standard distribution medium for ELM is one 5.25 inch 360K MS-DOS format diskette. ELM was developed in 1991. DEC, VAX, and VMS are trademarks of Digital Equipment Corporation. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation.

  6. The Influence of the Von Opel-Valier Experiments Upon German Rocket-Propelled Model Aircraft Development, 1920's-1930's

    Science.gov (United States)

    Winter, F. H.

    The following is a brief paper presented at the XXVIIIth International Astronautical Federation (IAF) Congress held in Prague, 25 September - 1 October 1977. Although it is generally assumed that recreational model rocketry, including commercially available kits, evolved during the Space Age, there were some interesting early predecessors, especially in Germany, that apparently arose as a result of the widespread reported rocket car and plane stunts of Fritz von Opel and Max Valier in the late 1920's.

  7. Computational survey of representative energetic materials as propellants for microthruster applications

    Science.gov (United States)

    Fuchs, Brian; Stec, Daniel, III

    2007-04-01

    Microthrusters are critical for the development of terrestrial micromissiles and nano air vehicles for reconnaissance, surveillance, and sensor emplacement. With the maturation of MEMS manufacturing technology, the physical components of the thrusters can be readily fabricated. The thruster type that is the most straightforward is chemical combustion of a propellant that is ignited by a heating element giving a single shot thrust. Arrays of MEMS manufactured thrusters can be ganged to give multiple firings. The basic model for such a system is a solid rocket motor. The desired elements for the propellant of a chemical thruster are high specific impulse (I sp), high temperature and pressure, and low molecular weight combustion gases. Since the combustion chamber of a microthruster is extremely small, the propellant material must be able to ignite, sustain and complete its burn inside the chamber. The propellant can be either a solid or a liquid. There are a large number of energetic materials available as candidates for a propellant for microthrusters. There has been no systematic evaluation of the available energetic materials as propellant candidates for microthrusters. This report summarizes computations done on a series of energetic materials to address their suitabilities as microthruster propellants.

  8. F. Gomez Arias' rocket vehicle project

    Science.gov (United States)

    Carreras, R.

    1977-01-01

    Research done by Spanish pioneer rocket scientists in the 19th century was investigated with major emphasis placed on F. Gomez Arias' rocket vehicle project. Arias, considered the world's first designer of rocket propelled, manned aircraft, was interested in solving the problem of space navigation. Major concerns included ascent and direction of heavier-than-airmachines, as well as ascent and direction of balloons.

  9. Test data from solid propellant plume aerodynamics test program in Ames 6 x 6 foot supersonic wind tunnel (shuttle test FA7) (Ames test 033-66)

    Science.gov (United States)

    Hair, L. M.

    1975-01-01

    The aerodynamic effects of plumes from hot combustion gases in the presence of a transonic external flow field were measured to advance plumes simulation technology, extend a previously acquired data base, and provide data to compare with the effects observed using cold gas plumes. A variety of underexpanded plumes issuing from the base of a strut-mounted ogive-cylinder body were produced by combusting solid propellant gas generators. The gas generator fired in a short-duration mode (200 to 300 msec). Propellants containing 16 percent and 2 percent A1 were used, with chamber pressures from 400 to 1800 psia. Conical nozzles of 15 deg half-angle were tested with area ratios of 4 and 8. Pressures were measured in the gas generator combustion chamber, along the nozzle wall, on the base, and along the body rear exterior. Schlieren photographs were taken for all tests. Test data are presented along with a description of the test setup and procedures.

  10. Ricardo Dyrgalla (1910-1970), pioneer of rocket development in Argentina

    Science.gov (United States)

    de León, Pablo

    2009-12-01

    One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

  11. Development status of M-V rocket structures and mechanisms

    Science.gov (United States)

    Onoda, Junjiro; Minesugi, Kenji; Watanabe, Naoyuki

    M-V is the next generation satellite launcher of the Mu rocket series of Institute of Space and Astronautical Science (ISAS). This paper describes the plan and the status of the development of its structure and mechanisms. The performance of the motor casings for the solid propellant rockets, which are the largest structural members in each stage, will be much improved by the introduction of new materials and new fabrication methods. For the nose-faring made of honeycomb sandwich shells with carbon fiber reinforced plastic (CFRP) face sheet, a new separation joint is being developed, which is composed of an expanding shielded mild detonating cord. A unique inter-stage joint between the 1st and 2nd stages is being developed in order to accommodate to the fire-in-the-hole (FITH) ignition of the 2nd stage rocket motor.

  12. Aluminum agglomeration involving the second mergence of agglomerates on the solid propellants burning surface: Experiments and modeling

    Science.gov (United States)

    Ao, Wen; Liu, Xin; Rezaiguia, Hichem; Liu, Huan; Wang, Zhixin; Liu, Peijin

    2017-07-01

    The agglomeration of aluminum particles usually occurs on the burning surface of aluminized composite propellants. It leads to low propellant combustion efficiency and high two-phase flow losses. To reach a thorough understanding of aluminum agglomeration behaviors, agglomeration processes, and particles size distribution of Al/AP/RDX/GAP propellants were studied by using a cinephotomicrography experimental technique, under 5 MPa. Accumulation, aggregation, and agglomeration phenomena of aluminum particles have been inspected, as well as the flame asymmetry of burning agglomerates. Results reveals that the dependency of the mean and the maximum agglomeration diameter to the burning rate and the virgin aluminum size have the same trend. A second-time mergence of multiple agglomerates on the burning surface is unveiled. Two typical modes of second mergence are concluded, based upon vertical and level movement of agglomerates, respectively. The latter mode is found to be dominant and sometimes a combination of the two modes may occur. A new model of aluminum agglomeration on the burning surface of composite propellants is derived to predict the particulates size distribution with a low computational amount. The basic idea is inspired from the well-known pocket models. The pocket size of the region formed by adjacent AP particles is obtained through scanning electron microscopy of the propellant cross-section coupled to an image processing method. The second mergence mechanism, as well as the effect of the burning rate on the agglomeration processes, are included in the present model. The mergence of two agglomerates is prescribed to occur only if their separation distance is less than a critical value. The agglomerates size distribution resulting from this original model match reasonably with the experimental data. Moreover, the present model gives superior results for mean agglomeration diameter compared to common empirical and pocket models. The average prediction

  13. Terminology and Assessment Methods of Solid Propellant Rocket Exhaust Signatures (Methodes d’Evaluation des Signatures des Propulseurs a Propergol Solide)

    Science.gov (United States)

    1993-02-01

    11 69D-80tJ0 Mu~nchcn 80 Spain Mr D,M. Barreiros Dr R. Dirscherl Departemento Energia y Deutsche Aerospace Propulsion Abteilung VAR 331 INTA Postfach... Solar scattering from less thtan a factor at I1) for smokte bmoke a Search lightl effert Near infrared S Glood S Poor P lrobably figir but relatively...rotational systems 8.01 to00o or solar aind earthishine regioni lredictiolls cai be made to a (1 Molecular Ilybrlde relatioii factor or 5 (I F. I [CI 1120) a

  14. Attitude Dynamics of a Spinning Rocket with Internal Fluid Whirling Motion

    Directory of Open Access Journals (Sweden)

    Marius Ionut MARMUREANU

    2014-06-01

    Full Text Available This paper evaluates the impact that helical motion of fluid products of combustion within the combustion chamber of a rocket can have on the attitude dynamics of rocket systems. By developing the study presented by Sookgaew (2004, we determined the configuration of the Coriolis moment components, which catch the impact of the combustion product’s whirling motion, for the radial and centripetal propellant burn pattern specific to S-5M and S-5K solid rocket motors. We continue the investigation of the effects of internal whirling motion of fluid products of combustion on the attitude behavior of variable mass systems of the rocket type by examining the spin motion and transverse attitude motion of such systems. The results obtained show that internal fluid whirling motion can cause appreciable deviations in spin rate predictions, and also affects the frequencies of the transverse angular velocity components.

  15. Magnesium Based Rockets for Martian Exploration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — We propose to develop Mg rockets for Martian ascent vehicle applications. The propellant can be acquired in-situ from MgO in the Martian regolith (5.1% Mg by mass)...

  16. The sky is falling III: The effect of deposition from static solid rocket motor tests on juvenile crops.

    Science.gov (United States)

    Doucette, William J; Curry, Eric; McNeill, Laurie S; Heavilin, Justin

    2017-12-01

    A mixture of combustion products (mainly hydrogen chloride, aluminum oxide, and water) and entrained soil, referred to as Test Fire Soil (TFS), can be deposited on crops during static solid rocket motor tests. The impact of a reported worst-case event was previously evaluated by exposing corn and alfalfa to 3200-gTFS/m(2) at 54days after emergence. Exposures via soil and leaves were evaluated separately. Reduced growth (soil exposure) and leaf "scorch" (leaf exposure) were attributed mainly to the high chloride concentrations in the TFS (56,000mg/kg). A follow-up study was conducted to evaluate the effect of a typical deposition event (70-gTFS/m(2), estimated by radar during several tests) and exposure (soil and leaves simultaneously) on juvenile corn, alfalfa, and winter wheat. Younger crops were used to examine potential age sensitivity differences. Impact was evaluated by comparing the growth, elemental composition, and leaf chlorophyll content of treated and untreated plants. The relationship between deposition exposure and response was also addressed. Growth of corn, alfalfa, and winter wheat exposed to a typical TFS loading was not impacted, although slightly elevated concentrations of aluminum and iron were found in the leaves. At the highest loadings used for the exposure-response experiment, concentrations of chloride and calcium were higher in TFS-exposed corn leaves than in the untreated leaves. Overall results indicate that exposure to a typical deposition event does not adversely impact juvenile crops and that younger plants may be less vulnerable to TFS. However, higher TFS loadings can cause leaf scorch and increase the leaf concentrations of some elements. Copyright © 2017 Elsevier B.V. All rights reserved.

  17. Laser rocket system analysis

    Science.gov (United States)

    Jones, W. S.; Forsyth, J. B.; Skratt, J. P.

    1979-01-01

    The laser rocket systems investigated in this study were for orbital transportation using space-based, ground-based and airborne laser transmitters. The propulsion unit of these systems utilizes a continuous wave (CW) laser beam focused into a thrust chamber which initiates a plasma in the hydrogen propellant, thus heating the propellant and providing thrust through a suitably designed nozzle and expansion skirt. The specific impulse is limited only by the ability to adequately cool the thruster and the amount of laser energy entering the engine. The results of the study showed that, with advanced technology, laser rocket systems with either a space- or ground-based laser transmitter could reduce the national budget allocated to space transportation by 10 to 345 billion dollars over a 10-year life cycle when compared to advanced chemical propulsion systems (LO2-LH2) of equal capability. The variation in savings depends upon the projected mission model.

  18. The thermal decomposition behavior of ammonium perchlorate and of an ammonium-perchlorate-based composite propellant

    Energy Technology Data Exchange (ETDEWEB)

    Behrens, R.; Minier, L.

    1998-03-24

    The thermal decomposition of ammonium perchlorate (AP) and ammonium-perchlorate-based composite propellants is studied using the simultaneous thermogravimetric modulated beam mass spectrometry (STMBMS) technique. The main objective of the present work is to evaluate whether the STMBMS can provide new data on these materials that will have sufficient detail on the reaction mechanisms and associated reaction kinetics to permit creation of a detailed model of the thermal decomposition process. Such a model is a necessary ingredient to engineering models of ignition and slow-cookoff for these AP-based composite propellants. Results show that the decomposition of pure AP is controlled by two processes. One occurs at lower temperatures (240 to 270 C), produces mainly H{sub 2}O, O{sub 2}, Cl{sub 2}, N{sub 2}O and HCl, and is shown to occur in the solid phase within the AP particles. 200{micro} diameter AP particles undergo 25% decomposition in the solid phase, whereas 20{micro} diameter AP particles undergo only 13% decomposition. The second process is dissociative sublimation of AP to NH{sub 3} + HClO{sub 4} followed by the decomposition of, and reaction between, these two products in the gas phase. The dissociative sublimation process occurs over the entire temperature range of AP decomposition, but only becomes dominant at temperatures above those for the solid-phase decomposition. AP-based composite propellants are used extensively in both small tactical rocket motors and large strategic rocket systems.

  19. On the combustion mechanisms of ZrH2 in double-base propellant.

    Science.gov (United States)

    Yang, Yanjing; Zhao, Fengqi; Yuan, Zhifeng; Wang, Ying; An, Ting; Chen, Xueli; Xuan, Chunlei; Zhang, Jiankan

    2017-12-13

    Metal hydrides are regarded as a series of promising hydrogen-supplying fuel for solid rocket propellants. Their effects on the energetic and combustion performances of propellants are closely related to their reaction mechanisms. Here we report a first attempt to determine the reaction mechanism of ZrH2, a high-density metal hydride, in the combustion of a double-base propellant to evaluate its potential as a fuel. ZrH2 is determined to possess good resistance to oxidation by nitrocellulose and nitroglycerine. Thus its combustion starts with dehydrogenation to generate H2 and metallic Zr. Subsequently, the newly formed Zr and H2 participate in the combustion and, especially, Zr melts and then combusts on the burning surface which favors the heat feedback to the propellant. This phenomenon is completely different from the combustion behavior of the traditional fuel Al, where the Al particles are ejected off the burning surface of the propellant to get into the luminous flame zone to burn. The findings in this work validate the potential of ZrH2 as a hydrogen-supplying fuel for double-base propellants.

  20. Mechanical and Combustion Performance of Multi-Walled Carbon Nanotubes as an Additive to Paraffin-Based Solid Fuels for Hybrid Rockets

    Science.gov (United States)

    Larson, Daniel B.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth, K.; Koo, Joseph H.; Story, George

    2012-01-01

    Paraffin-based solid fuels for hybrid rocket motor applications are recognized as a fastburning alternative to other fuel binders such as HTPB, but efforts to further improve the burning rate and mechanical properties of paraffin are still necessary. One approach that is considered in this study is to use multi-walled carbon nanotubes (MWNT) as an additive to paraffin wax. Carbon nanotubes provide increased electrical and thermal conductivity to the solid-fuel grains to which they are added, which can improve the mass burning rate. Furthermore, the addition of ultra-fine aluminum particles to the paraffin/MWNT fuel grains can enhance regression rate of the solid fuel and the density impulse of the hybrid rocket. The multi-walled carbon nanotubes also present the possibility of greatly improving the mechanical properties (e.g., tensile strength) of the paraffin-based solid-fuel grains. For casting these solid-fuel grains, various percentages of MWNT and aluminum particles will be added to the paraffin wax. Previous work has been published about the dispersion and mixing of carbon nanotubes.1 Another manufacturing method has been used for mixing the MWNT with a phenolic resin for ablative applications, and the manufacturing and mixing processes are well-documented in the literature.2 The cost of MWNT is a small fraction of single-walled nanotubes. This is a scale-up advantage as future applications and projects will require low cost additives to maintain cost effectiveness. Testing of the solid-fuel grains will be conducted in several steps. Dog bone samples will be cast and prepared for tensile testing. The fuel samples will also be analyzed using thermogravimetric analysis and a high-resolution scanning electron microscope (SEM). The SEM will allow for examination of the solid fuel grain for uniformity and consistency. The paraffin-based fuel grains will also be tested using two hybrid rocket test motors located at the Pennsylvania State University s High Pressure

  1. Comparison of vibrations of a combination of solid-rocket launch vehicle and payload during a ground firing and launching

    Science.gov (United States)

    Schoenster, J. A.; Pierce, H. B.

    1975-01-01

    The results of a study into the environmental vibrations of a payload mounted on the Nike rocket launch vehicle were presented. Data were obtained during the flight acceptance test of the payload, the firing of the total vehicle in a special test stand, and the powered and unpowered flights of the vehicle. The vibrational response of the structure was measured. Data were also obtained on the fluctuating pressure on the outside surface of the vehicle and inside the forward and after ends of the rocket chamber. A comparison of the data from the three test conditions indicated that external pressure fluctuations were the major source of vibrations in the payload area, and pressure fluctuations within the rocket motor were the major source of vibrations contiguous to the payload area.

  2. An evaluation of the total quality management implementation strategy for the advanced solid rocket motor project at NASA's Marshall Space Flight Center. M.S. Thesis - Tennessee Univ.

    Science.gov (United States)

    Schramm, Harry F.; Sullivan, Kenneth W.

    1991-01-01

    An evaluation of the NASA's Marshall Space Flight Center (MSFC) strategy to implement Total Quality Management (TQM) in the Advanced Solid Rocket Motor (ASRM) Project is presented. The evaluation of the implementation strategy reflected the Civil Service personnel perspective at the project level. The external and internal environments at MSFC were analyzed for their effects on the ASRM TQM strategy. Organizational forms, cultures, management systems, problem solving techniques, and training were assessed for their influence on the implementation strategy. The influence of ASRM's effort was assessed relative to its impact on mature projects as well as future projects at MSFC.

  3. Regenerative cooling for liquid rocket engines

    Science.gov (United States)

    Qi, Feng

    1995-01-01

    Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocket engines. Regenerative cooling is an advanced method which can ensure not only the proper running but also higher performance of a rocket engine. The theoretical model is complicated, it relates to fluid dynamics, heat transfer, combustion, etc... In this paper, a regenerative cooling model is presented. Effects such as radiation, heat transfer to environment, variable thermal properties and coking are included in the model. This model can be applied to all kinds of liquid propellant rocket engines as well as similar constructions. The modularized computer code is completed in the work.

  4. Safety and Performance Advantages of Nitrous Oxide Fuel Blends (NOFBX) Propellants for Manned and Unmanned Spaceflight Applications

    Science.gov (United States)

    Taylor, R.

    2012-01-01

    Hydrazine, N2H4, is the current workhorse monopropellant in the spacecraft industry. Although widely used since the 1960's, hydrazine is highly toxic and its specific impulse (ISP) performance of ~230s is far lower than bipropellants and solid motors. NOFBX™ monopropellants were originally developed under NASA's Mars Advanced Technology program (2004-2007) for deep space Mars missions. This work focused on characterizing various Nitrous Oxide Fuel Blend (NOFB) monopropellants which exhibited many favorable attributes to include: (1) Mono-propulsion, (2) Isp > 320s, (3) Non-toxic constituents, (4) Non-toxic effluents, (5) Low Cost, (6) High Density Specific Impulse, (7) Non-cryogenic, (8) Wide Storable Temperature Range, (9) Deeply throttlable [between 5 - 100lbs], (10) Self Pressurizing, (11) Wide Range of materials compatibility, along with many, many other benefits. All rocket propellants carry with them a history or stigma associated with either the development or implementation of that propellant and NOFBX™ is no exception. This paper examines the benefits of NOFBX™ propellants while addressing or dispelling a number of critiques N2O based propellants acquired through the decades of rocket propellant testing.

  5. How Does Rocket Propulsion Work? The most common answer to ...

    Indian Academy of Sciences (India)

    The most common answer to the above question is – hot jet of gas comes out of the nozzle of the rocket engine at high speeds and as a reaction the rocket moves (is propelled) in the opposite direction [1]. But is this answer right? Let us explore what goes on inside a rocket engine and arrive at the right answer. Generally ...

  6. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Science.gov (United States)

    Jones, Carl P.; Robertson, Elizabeth H.; Koelbl, Mary Beth; Singer, Chris

    2016-01-01

    Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

  7. Rocket Science at the Nanoscale.

    Science.gov (United States)

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

    Autonomous propulsion at the nanoscale represents one of the most challenging and demanding goals in nanotechnology. Over the past decade, numerous important advances in nanotechnology and material science have contributed to the creation of powerful self-propelled micro/nanomotors. In particular, micro- and nanoscale rockets (MNRs) offer impressive capabilities, including remarkable speeds, large cargo-towing forces, precise motion controls, and dynamic self-assembly, which have paved the way for designing multifunctional and intelligent nanoscale machines. These multipurpose nanoscale shuttles can propel and function in complex real-life media, actively transporting and releasing therapeutic payloads and remediation agents for diverse biomedical and environmental applications. This review discusses the challenges of designing efficient MNRs and presents an overview of their propulsion behavior, fabrication methods, potential rocket fuels, navigation strategies, practical applications, and the future prospects of rocket science and technology at the nanoscale.

  8. Low-thrust rocket trajectories

    Energy Technology Data Exchange (ETDEWEB)

    Keaton, P.W.

    1987-03-01

    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report.

  9. Low-thrust rocket trajectories

    Energy Technology Data Exchange (ETDEWEB)

    Keaton, P.W.

    1986-01-01

    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report. 57 refs., 10 figs.

  10. Measured Temperautres of Solid Rocket Motors Dump Stored in the Tropics and Desert. Part 3. Desert Storage

    Science.gov (United States)

    1977-05-01

    pubUcation UNCLASSIFIED.) 9 . Storage Temperature of Explotive Hazard Magazines. Part 5. Carrihbean and Mid-Atlantic, by I. S. Kurotori and H... fired hardware was much more plentiful. However, to utilize the once- fired hardware, an inexpensive durable propellant simulant was needed. The...also exposed because it was found necessary, during the Viet Nam conflict, to remove the packaging from the rounds in order to be ready for a fire

  11. Hybrid rocket engine research program at Ryerson University

    Energy Technology Data Exchange (ETDEWEB)

    Karpynczyk, J.; Greatrix, D.R. [Ryerson Polytechnic Univ., Toronto, ON (Canada). Dept. of Aerospace Engineering

    2007-07-01

    Hybrid rocket engines (HREs) are a combination of solid and liquid propellant rocket engine designs. A solid fuel grain is located in the main combustion chamber and nozzle aft, while a stored liquid or gaseous oxidizer source supplies the required oxygen content through a throttle valve, for combustion downstream in the main chamber. HREs have drawn significant interest in certain flight applications, as they can be advantageous in terms of cost, ease and safety in storage, controllability in flight, and availability of propellant constituents. Key factors that will lead to further practical usage of HREs for flight applications are their predictability and reproducibility of operational performance. This paper presented information on studies being conducted at Ryerson University aimed at analyzing and testing the performance of HREs. It discussed and illustrated the conventional HRE and analyzed engine performance considerations such as the fuel regression rate, mass flux about the fuel surface, burning rate, and zero transformation parameter. Other factors relating to HRE performance that were presented included induced forward and aft oxidizer flow swirl effects as a means for augmenting the fuel regression rate, stoichiometric grain length issues, and feed system stability. Last, the paper presented a simplified schematic diagram of a proposed thrust/test stand for HRE test firings. 2 refs., 3 figs.

  12. Enhanced alkaline hydrolysis and biodegradability studies of nitrocellulose-bearing missile propellant

    Science.gov (United States)

    Sidhoum, Mohammed; Christodoulatos, Christos; Su, Tsan-Liang; Redis, Mercurios

    1995-01-01

    Large amounts of energetic materials which have been accumulated over the years in various manufacturing and military installations must be disposed of in an environmentally sound manner. Historically, the method of choice for destruction of obsolete or aging energetic materials has been open burning or open detonation (OB/OD). This destruction approach has become undesirable due to air pollution problems. Therefore, there is a need for new technologies which will effectively and economically deal with the disposal of energetic materials. Along those lines, we have investigated a chemical/biological process for the safe destruction and disposal of a double base solid rocket propellant (AHH), which was used in several 8 inch projectile systems. The solid propellant is made of nitrocellulose and nitroglycerin as energetic components, two lead salts which act as ballistic modifiers, triacetin as a plasticizer and 2-Nitrodiphenylamine (2-NDPA) as a stabilizer. A process train is being developed to convert the organic components of the propellant to biodegradable products and remove the lead from the process stream. The solid propellant is first hydrolyzed through an enhanced alkaline hydrolysis process step. Following lead removal and neutralization, the digested liquor rich in nitrates and nitrites is found to be easily biodegradable. The digestion rate of the intact ground propellant as well as the release of nitrite and nitrate groups were substantially increased when ultrasound were supplied to the alkaline reaction medium compared to the conventional alkaline hydrolysis. The effects of reaction time, temperature, sodium hydroxide concentration and other relevant parameters on the digestion efficiency and biodegradability have been studied. The present work indicates that the AHH propellant can be disposed of safely with a combination of physiochemical and biological processes.

  13. Torpedo Rockets

    Science.gov (United States)

    2004-01-01

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  14. Rocket Flight.

    Science.gov (United States)

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

    Describes an activity for designing, building, and launching rockets that provides students with an intrinsically motivating and real-life application of what could have been classroom-only concepts. Includes rocket design guidelines and a sample grading rubric. (KHR)

  15. Transient combustion in hybrid rockets

    Science.gov (United States)

    Karabeyoglu, Mustafa Arif

    1998-09-01

    Hybrid rockets regained interest recently as an alternative chemical propulsion system due to their advantages over the solid and liquid systems that are currently in use. Development efforts on hybrids revealed two important problem areas: (1) low frequency instabilities and (2) slow transient response. Both of these are closely related to the transient behavior which is a poorly understood aspect of hybrid operation. This thesis is mainly involved with a theoretical study of transient combustion in hybrid rockets. We follow the methodology of identifying and modeling the subsystems of the motor such as the thermal lags in the solid, boundary layer combustion and chamber gasdynamics from a dynamic point of view. We begin with the thermal lag in the solid which yield the regression rate for any given wall heat flux variation. Interesting phenomena such as overshooting during throttling and the amplification and phase lead regions in the frequency domain are discovered. Later we develop a quasi-steady transient hybrid combustion model supported with time delays for the boundary layer processes. This is integrated with the thermal lag system to obtain the thermal combustion (TC) coupled response. The TC coupled system with positive delays generated low frequency instabilities. The scaling of the instabilities are in good agreement with actual motor test data. Finally, we formulate a gasdynamic model for the hybrid chamber which successfully resolves the filling/emptying and longitudinal acoustic behavior of the motor. The TC coupled system is later integrated to the gasdynamic model to obtain the overall response (TCG coupled system) of gaseous oxidizer motors with stiff feed systems. Low frequency instabilities were also encountered for the TCG coupled system. Apart from the transient investigations, the regression rate behavior of liquefying hybrid propellants such as solid cryogenic materials are also studied. The theory is based on the possibility of enhancement

  16. Particle size reduction of propellants by cryocycling

    Energy Technology Data Exchange (ETDEWEB)

    Whinnery, L.; Griffiths, S.; Lipkin, J. [and others

    1995-05-01

    Repeated exposure of a propellant to liquid nitrogen causes thermal stress gradients within the material resulting in cracking and particle size reduction. This process is termed cryocycling. The authors conducted a feasibility study, combining experiments on both inert and live propellants with three modeling approaches. These models provided optimized cycle times, predicted ultimate particle size, and allowed crack behavior to be explored. Process safety evaluations conducted separately indicated that cryocycling does not increase the sensitivity of the propellants examined. The results of this study suggest that cryocycling is a promising technology for the demilitarization of tactical rocket motors.

  17. Air-Breathing Rocket Engines

    Science.gov (United States)

    1998-01-01

    This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  18. Combustion characteristics of SMX and SMX based propellants

    Science.gov (United States)

    Reese, David A.

    This work investigates the combustion of the new solid nitrate ester 2,3-hydroxymethyl-2,3-dinitro-1,4-butanediol tetranitrate (SMX, C6H 8N6O16). SMX was synthesized for the first time in 2008. It has a melting point of 85 °C and oxygen balance of 0% to CO 2, allowing it to be used as an energetic additive or oxidizer in solid propellants. In addition to its neat combustion characteristics, this work also explores the use of SMX as a potential replacement for nitroglycerin (NG) in double base gun propellants and as a replacement for ammonium perchlorate in composite rocket propellants. The physical properties, sensitivity characteristics, and combustion behaviors of neat SMX were investigated. Its combustion is stable at pressures of up to at least 27.5 MPa (n = 0.81). The observed flame structure is nearly identical to that of other double base propellant ingredients, with a primary flame attached at the surface, a thick isothermal dark zone, and a luminous secondary flame wherein final recombination reactions occur. As a result, the burning rate and primary flame structure can be modeled using existing one-dimensional steady state techniques. A zero gas-phase activation energy approximation results in a good fit between modeled and observed behavior. Additionally, SMX was considered as a replacement for nitroglycerin in a double base propellant. Thermochemical calculations indicate improved performance when compared with the common double base propellant JA2 at SMX loadings above 40 wt-%. Also, since SMX is a room temperature solid, migration may be avoided. Like other nitrate esters, SMX is susceptible to decomposition over long-term storage due to the presence of excess acid in the crystals; the addition of stabilizers (e.g., derivatives of urea) during synthesis should be sufficient to prevent this. the addition of Both unplasticized and plasticized propellants were formulated. Thermal analysis of unplasticized propellant showed a distinct melt

  19. Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

    Science.gov (United States)

    Larson, Daniel B.; Desain, John D.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth K.; Borduin, Russell; Koo, Joseph H.; Brady, Brian B.; Curtiss, Thomas J.; Story, George

    2012-01-01

    This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

  20. Combustion efficiency and pyrochemical properties of micron-sized metal particles as the components of modified double-base propellant

    Science.gov (United States)

    Wu, Xiong-Gang; Yan, Qi-Long; Guo, Xin; Qi, Xiao-Fei; Li, Xiao-Jiang; Wang, Ke-Qiang

    2011-04-01

    The combustion efficiency of metallized propellants are investigated and compared to the corresponding blank propellant in order to evaluate the actual effect of the metals in solid rocket applications. The image analysis coupled with energy dispersive spectrometry (EDS) analysis on a scanning electron microscope (SEM) and quantitative X-ray diffraction analysis were applied to the characterization of the original metal particles such as aluminum (Al), magnesium (Mg), boron (B), nickel (Ni), and Mg-Al alloy (Mg/Al) and their condensed combustion products. Under the explored operating conditions, the results confirm that the metallized propellants show heterogeneous diffusing flame, with significant change in pressure sensitivity, and larger aggregation/agglomeration phenomena in combustion products than that of the blank propellant. Besides, the chemical reactions in condensed phase and gas phase which control the burning process and combustion efficiency of the double-base propellant containing different metal particles were systematically investigated and descriptions of the detailed reaction mechanisms from solid phase to liquid phase or to gas phase are also included. It was indicated that the combustion efficiency is favored by the activity and melting points of the metals.

  1. Particle Size Distributions Measured in the Stratospheric Plumes of Three Rockets During the ACCENT Missions

    Science.gov (United States)

    Wiedinmyer, C.; Brock, C. A.; Reeves, J. M.; Ross, M. N.; Schmid, O.; Toohey, D.; Wilson, J. C.

    2001-12-01

    The global impact of particles emitted by rocket engines on stratospheric ozone is not well understood, mainly due to the lack of comprehensive in situ measurements of the size distributions of these emitted particles. During the Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) missions in 1999, the NASA WB-57F aircraft carried the University of Denver N-MASS and FCAS instruments into the stratospheric plumes from three rockets. Size distributions of particles with diameters from 4 to approximately 2000 nm were calculated from the instrument measurements using numerical inversion techniques. The data have been averaged over 30-second intervals. The particle size distributions observed in all of the rocket plumes included a dominant mode near 60 nm diameter, probably composed of alumina particles. A smaller mode at approximately 25 nm, possibly composed of soot particles, was seen in only the plumes of rockets that used liquid oxygen and kerosene as a propellant. Aircraft exhaust emitted by the WB-57F was also sampled; the size distributions within these plumes are consistent with prior measurements in aircraft plumes. The size distributions for all rocket intercepts have been fitted to bimodal, lognormal distributions to provide input for global models of the stratosphere. Our data suggest that previous estimates of the solid rocket motor alumina size distributions may underestimate the alumina surface area emission index, and so underestimate the particle surface area available for heterogeneous chlorine activation reactions in the global stratosphere.

  2. Mars McLOX Rocket Propulsion System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Methane and Carbon Monoxide/LOX rocket (MCLOX) is a technology for accomplishing ascent from Mars. Current Mars in-situ propellant production (ISPP) technologies...

  3. Ultraviolet photographic pyrometer used in rocket exhaust analysis

    Science.gov (United States)

    Levin, B. P.

    1966-01-01

    Ultraviolet photographic pyrometer investigates the role of carbon as a thermal radiator and determines the geometry, location, and progress of afterburning phenomena in the exhaust plume of rocket engines using liquid oxygen/RP-1 as propellant.

  4. Performance and Cost Evaluation of Cryogenic Solid Propulsion Systems

    Science.gov (United States)

    Adirim, Harry; Lo, Roger; Knecht, Thomas; Reinbold, Georg-Friedrich; Poller, Sascha

    2002-01-01

    Under the sponsorship of the German Aerospace Center DLR, Cryogenic Solid Propulsion (CSP) is now in its 6th year of R&D. The development proceeds as a joint international university-, small business-, space industry- and professional research effort (Berlin University of Technology / AI: Aerospace Institute, Berlin / Bauman Moscow State Technical University, Russia / ASTRIUM GmbH, Bremen / Fraunhofer Institute for Chemical Technology, Berghausen). This paper aims at introducing CSP as a novel type of chemical propellant that uses frozen liquids as Oxygen (SOX) or Hydrogen Peroxide (SH2O2) inside of a coherent solid Hydrocarbon (PE, PU or HTPB) matrix in solid rocket motors. Theoretically any conceivable chemical rocket propellant combination (including any environmentally benign ,,green propellant") can be used in solid rocket propellant motors if the definition of solids is not restricted to "solid at ambient temperature". The CSP concept includes all suitable high energy propellant combinations, but is not limited to them. Any liquid or hybrid bipropellant combination is (Isp-wise) superior to any conventional solid propellant formulation. While CSPs do share some of the disadvantages of solid propulsion (e.g. lack of cooling fluid and preset thrust-time function), they definitely share one of their most attractive advantages: the low number of components that is the base for high reliability and low cost of structures. In this respect, CSPs are superior to liquid propellant rocket motors with whom, they share the high Isp performance. High performance, low cost, low pollution CSP technology could bring about a near term improvement for chemical Earth-to-orbit high thrust propulsion. In the long run it could surpass conventional chemical propulsion because it is better suited for applying High Energy Density Matter (HEDM) than any other mode of propulsion. So far, ongoing preliminary analyses have not shown any insuperable problems in areas of concern, such as

  5. Tip-modified Propellers

    DEFF Research Database (Denmark)

    Andersen, Poul

    1999-01-01

    wings. The literature on four different designs is reviewed: the end-plate propeller; the two-sided, shifted end-plate propeller; the tip-fin propeller; and the bladelet propeller. The conclusion is that it is indeed possible to design tip-modified propellers that, relative to an optimum conventional...

  6. Thermal Hydraulics Design and Analysis Methodology for a Solid-Core Nuclear Thermal Rocket Engine Thrust Chamber

    Science.gov (United States)

    Wang, Ten-See; Canabal, Francisco; Chen, Yen-Sen; Cheng, Gary; Ito, Yasushi

    2013-01-01

    Nuclear thermal propulsion is a leading candidate for in-space propulsion for human Mars missions. This chapter describes a thermal hydraulics design and analysis methodology developed at the NASA Marshall Space Flight Center, in support of the nuclear thermal propulsion development effort. The objective of this campaign is to bridge the design methods in the Rover/NERVA era, with a modern computational fluid dynamics and heat transfer methodology, to predict thermal, fluid, and hydrogen environments of a hypothetical solid-core, nuclear thermal engine the Small Engine, designed in the 1960s. The computational methodology is based on an unstructured-grid, pressure-based, all speeds, chemically reacting, computational fluid dynamics and heat transfer platform, while formulations of flow and heat transfer through porous and solid media were implemented to describe those of hydrogen flow channels inside the solid24 core. Design analyses of a single flow element and the entire solid-core thrust chamber of the Small Engine were performed and the results are presented herein

  7. Two-step rocket engine bipropellant valve concept

    Science.gov (United States)

    Capps, J. E.; Ferguson, R. E.; Pohl, H. O.

    1969-01-01

    Initiating combustion of altitude control rocket engines in a precombustion chamber of ductile material reduces high pressure surges generated by hypergolic propellants. Two-step bipropellant valve concepts control initial propellant flow into precombustion chamber and subsequent full flow into main chamber.

  8. Analysis of a Radioisotope Thermal Rocket Engine

    Science.gov (United States)

    Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.

    2017-01-01

    The Triton Hopper is a concept for a vehicle to explore the surface of Neptunes moon Triton, which uses a radioisotope heated rocket engine and in-situ propellant acquisition. The initial Triton Hopper conceptual design stores pressurized Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through a thermal block during engine operation, as opposed to storing gas at a high temperature.

  9. Propellant Nonlinear Constitutive Theory Extension: Preliminary Results.

    Science.gov (United States)

    1983-08-01

    r) dr f1 dr 0 *where: u : stress e: strain E(t) = relaxation modulus ( PRONY series representation using a matrix solution for curve fitting data...series of unaxial tests were conducted with two materials - a PEAN and an HTPB propellant. The data was fit to each of the nonlinear theories.* Then the...PRELIMINARY STUDY Generally speaking, solid propellants may be considered as lightly cross- linked long-chain polymers , highly filled with coarse solid

  10. ADN – The new oxidizer around the corner for an environmentally friendly smokeless propellant

    Directory of Open Access Journals (Sweden)

    Márcio Y. Nagamachi

    2009-06-01

    Full Text Available The search for a smokeless propellant has encouraged scientists and engineers to look for a chlorine-free oxidizer as a substitute for AP (ammonium perchlorate. Endeavors seemed to come to an end when ADN (ammonium dinitramide appeared in the West in the early 1990s. Although some drawbacks soon became apparent by that time, the foremost obstacle for its use in rocket-motors came from the patent originally applied for in the United States in 1990. Furthermore, environmental concerns have also increased during these two decades. Ammonium perchlorate is believed to cause thyroid cancer by contaminating soil and water. In addition, AP produces hydrogen chloride during burning which can cause acid rain and ozone layer depletion. Unlike AP, ADN stands for both smokeless and green propellant. Since then, much progress has been made in its development in synthesis, re-shaping, microencapsulation and solid propellant. The high solubility of ADN in water has also allowed its application as liquid monopropellant. Tests have revealed Isp (specific impulse superior to that normally observed with hydrazine, one of the most harmful and hazardous liquid propellants. With constraints of use, along with the patent near to expiry, scientists and engineers are rushing to complete developments and patents until then.

  11. Theoretical and Experimental Analysis of the Physics of Water Rockets

    Science.gov (United States)

    Barrio-Perotti, R.; Blanco-Marigorta, E.; Fernandez-Francos, J.; Galdo-Vega, M.

    2010-01-01

    A simple rocket can be made using a plastic bottle filled with a volume of water and pressurized air. When opened, the air pressure pushes the water out of the bottle. This causes an increase in the bottle momentum so that it can be propelled to fairly long distances or heights. Water rockets are widely used as an educational activity, and several…

  12. Low thrust chemical rocket technology

    Science.gov (United States)

    Schneider, Steven J.

    1992-01-01

    An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher

  13. Green Propellant Demonstration with Hydrazine Catalyst of F-16 Emergency Power Unit

    Science.gov (United States)

    Robinson, Joel W.; Brechbill, Shawn

    2015-01-01

    Some space vehicle and aircraft Auxiliary Power Units (APUs) use hydrazine propellant for generating power. Hydrazine is a toxic, hazardous fuel which requires special safety equipment and processes for handling and loading. In recent years, there has been development of two green propellants that could enable their use in APU's: the Swedish LMP-103S and the Air Force Research Laboratory (AFRL) AF-M315E. While there has been work on development of these propellants for thruster applications (Prisma and Green Propulsion Infusion Mission, respectively), there has been less focus on the application to power units. Beginning in 2012, an effort was started by the Marshall Space Flight Center (MSFC) on the APU application. The MSFC plan was to demonstrate green propellants with residual Space Shuttle hardware. The principal investigator was able to acquire a Solid Rocket Booster gas generator and an Orbiter APU. Since these test assets were limited in number, an Air Force equivalent asset was identified: the F-16 Emergency Power Unit (EPU). In June 2013, two EPU's were acquired from retired aircraft located at Davis Monthan Air Force Base. A gas generator from one of these EPU's was taken out of an assembly and configured for testing with a version of the USAF propellant with a higher water content (AF-M315EM) to reduce decomposition temperatures. Testing in November 2014 has shown that this green propellant is reactive with the Hydrazine catalyst (Shell 405) generating 300 psi of pressure with the existing F-16 EPU configuration. This paper will highlight the results of MSFC testing in collaboration with AFRL.

  14. Optimization of Tape Winding Process Parameters to Enhance the Performance of Solid Rocket Nozzle Throat Back Up Liners using Taguchi's Robust Design Methodology

    Science.gov (United States)

    Nath, Nayani Kishore

    2017-08-01

    The throat back up liners is used to protect the nozzle structural members from the severe thermal environment in solid rocket nozzles. The throat back up liners is made with E-glass phenolic prepregs by tape winding process. The objective of this work is to demonstrate the optimization of process parameters of tape winding process to achieve better insulative resistance using Taguchi's robust design methodology. In this method four control factors machine speed, roller pressure, tape tension, tape temperature that were investigated for the tape winding process. The presented work was to study the cogency and acceptability of Taguchi's methodology in manufacturing of throat back up liners. The quality characteristic identified was Back wall temperature. Experiments carried out using L 9 ' (34) orthogonal array with three levels of four different control factors. The test results were analyzed using smaller the better criteria for Signal to Noise ratio in order to optimize the process. The experimental results were analyzed conformed and successfully used to achieve the minimum back wall temperature of the throat back up liners. The enhancement in performance of the throat back up liners was observed by carrying out the oxy-acetylene tests. The influence of back wall temperature on the performance of throat back up liners was verified by ground firing test.

  15. Insights into the Behavior of Potential Structural Failures Originating from Localized High Stress Regions in Configurations Relevant to Solid Rocket Motor Nozzles

    Science.gov (United States)

    McCutcheon, David Matthew

    2017-01-01

    During the structural certification effort for the Space Launch System solid rocket booster nozzle, it was identified that no consistent method for addressing local negative margins of safety in non-metallic materials had been developed. Relevant areas included bond-line terminations and geometric features in the composite nozzle liners. In order to gain understanding, analog test specimens were designed that very closely mimic the conditions in the actual full scale hardware. Different locations in the nozzle were represented by different analog specimen designs. This paper describes those tests and corresponding results. Finite element analysis results for the tests are presented. Strain gage correlation of the analysis to the test results is addressed. Furthermore, finite fracture mechanics (a coupled stress and energy failure criterion) is utilized to predict the observed crack pop-in loads for the different configurations. The finite fracture mechanics predictions are found to be within a 10% error relative to the average measured pop-in load for each of four configurations. Initiation locations, arrest behaviors, and resistances to further post-arrest crack propagation are also discussed.

  16. Development of Flow and Heat Transfer Models for the Carbon Fiber Rope in Nozzle Joints of the Space Shuttle Reusable Solid Rocket Motor

    Science.gov (United States)

    Wang, Q.; Ewing, M. E.; Mathias, E. C.; Heman, J.; Smith, C.; McCool, Alex (Technical Monitor)

    2001-01-01

    Methodologies have been developed for modeling both gas dynamics and heat transfer inside the carbon fiber rope (CFR) for applications in the space shuttle reusable solid rocket motor joints. Specifically, the CFR is modeled using an equivalent rectangular duct with a cross-section area, friction factor and heat transfer coefficient such that this duct has the same amount of mass flow rate, pressure drop, and heat transfer rate as the CFR. An equation for the friction factor is derived based on the Darcy-Forschheimer law and the heat transfer coefficient is obtained from pipe flow correlations. The pressure, temperature and velocity of the gas inside the CFR are calculated using the one-dimensional Navier-Stokes equations. Various subscale tests, both cold flow and hot flow, have been carried out to validate and refine this CFR model. In particular, the following three types of testing were used: (1) cold flow in a RSRM nozzle-to-case joint geometry, (2) cold flow in a RSRM nozzle joint No. 2 geometry, and (3) hot flow in a RSRM nozzle joint environment simulator. The predicted pressure and temperature history are compared with experimental measurements. The effects of various input parameters for the model are discussed in detail.

  17. Ramjet Application Possibilities for Increasing Fire Range of the Multiple Launch Rocket Systems Ammunition

    Directory of Open Access Journals (Sweden)

    V. N. Zubov

    2015-01-01

    Full Text Available The article considers a possibility to increase a flying range of the perspective rockets equipped with the control unit with aerodynamic controllers for the multiple launch rocket systems “Smerch”.To increase a flying range and reduce a starting mass of the rocket, the paper studies a possibility to replace the single-mode rocket engine used in the solid-fuel rocket motor for the direct-flow propulsion jet engine (DFPJE with not head sector air intakes. The DFPJE is implemented according to the classical scheme with a fuel charged in the combustion chamber. A separated solid propellant starting accelerator provides the rocket acceleration to reach a speed necessary for the DFPJE to run.When designing the DFPJE a proper choice of not head air intake parameters is one of the most difficult points. For this purpose a COSMOS Flow Simulation software package and analytical dependences were used to define the following: a boundary layer thickness where an air intake is set, maximum permissible and appropriate angles of attack and deviation angles of controllers at the section where the DFPJE works, and some other parameters as well.Calculation of DFPJE characteristics consisted in determining parameters of an air-gas path of the propulsion system, geometrical sizes of the pipeline flow area, sizes of a fuel charge, and dependence of the propulsion system impulse on the flight height and speed. Calculations were performed both in thermodynamic statement of problem and in using software package of COSMOS Flow Simulation.As a result of calculations and design engineering activities the air intake profile is created and mass-dimensional characteristics of DFPJE are defined. Besides, calculations of the starting solid fuel accelerator were carried out. Further design allowed us to create the rocket shape, estimate its mass-dimensional characteristics, and perform ballistic calculations, which proved that achieving a range of 120 km for the rocket is

  18. Mars Ascent Vehicle-Propellant Aging

    Science.gov (United States)

    Dankanich, John; Rousseau, Jeremy; Williams, Jacob

    2015-01-01

    This project is to develop and test a new propellant formulation specifically for the Mars Ascent Vehicle (MAV) for the robotic Mars Sample Return mission. The project was initiated under the Planetary Sciences Division In-Space Propulsion Technology (ISPT) program and is continuing under the Mars Exploration Program. The two-stage, solid motor-based MAV has been the leading MAV solution for more than a decade. Additional studies show promise for alternative technologies including hybrid and bipropellant options, but the solid motor design has significant propellant density advantages well suited for physical constraints imposed while using the SkyCrane descent stage. The solid motor concept has lower specific impulse (Isp) than alternatives, but if the first stage and payload remain sufficiently small, the two-stage solid MAV represents a potential low risk approach to meet the mission needs. As the need date for the MAV slips, opportunities exist to advance technology with high on-ramp potential. The baseline propellant for the MAV is currently the carboxyl terminated polybutadiene (CTPB) based formulation TP-H-3062 due to its advantageous low temperature mechanical properties and flight heritage. However, the flight heritage is limited and outside the environments, the MAV must endure. The ISPT program competed a propellant formulation project with industry and selected ATK to develop a new propellant formulation specifically for the MAV application. Working with ATK, a large number of propellant formulations were assessed to either increase performance of a CTPB propellant or improve the low temperature mechanical properties of a hydroxyl terminated polybutadiene (HTPB) propellant. Both propellants demonstrated potential to increase performance over heritage options, but an HTPB propellant formulation, TP-H-3544, was selected for production and testing. The test plan includes propellant aging first at high vacuum conditions, representative of the Mars transit

  19. Focused Rocket-Ejector RBCC Experiments

    Science.gov (United States)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    This document reports the results of additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Perm State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3rd generation Reusable Launch Vehicles (RLV). The two tasks conducted under this program build on earlier NASA MSFC funded research program on rocket ejector investigations. The first task continued a systematic investigation of the improvements provided by a gaseous hydrogen (GHz)/oxygen (GO2) twin thruster RBCC rocket ejector system over a single rocket system. In a similar vein, the second task continued investigations into the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. For the GH2/GO2 propellant rocket ejector experiments, high frequency measurements of the pressure field within the system were also made to understand the unsteady behavior of the flowfield.

  20. Perancangan Propeler Self-Propelled Barge

    Directory of Open Access Journals (Sweden)

    Billy Teguh kurniawan

    2013-03-01

    Full Text Available Makalah ini menyampaikan suatu penelitian tentang perancangan propeler yang optimal beserta pemilihan daya mesin yang efisien pada self-propelled barge dengan memperhitungkan besarnya nilai tahanan dari barge tersebut. Dengan penambahan sistem propulsi, diharapkan barge dapat beroperasi dengan lebih efisien dibandingkan saat barge beroperasi menggunakan sistem towing atau ditarik tug boat. Perhitungan tahanan barge dilakukan menggunakan metode Holtrop dan Guldhammer-Harvald sehingga dapat diperhi-tungkan geometri dan jenis propeler yang optimal beserta daya mesin yang efisien untuk barge. Propeler yang dianalisis adalah propeler tipe B-Troost Series, sedangkan variasi yang dilakukan untuk perencanaan propeler pada kajian ini adalah variasi putaran propeler pada rentang antara 310-800 rpm, serta variasi jumlah daun pada rentang tiga, empat, lima, dan enam. Besarnya nilai tahanan self-propelled barge untuk metode Holtrop adalah 105.91 kilonewton, sedangkan hasil per-hitungan dari metode Guldhammer-Harvald didapatkan nilai sebesar 109.14 kilonewton. Tipe propeler yang dipilih setelah dilakukan uji kavitasi adalah tipe Troost Series B4-40, dengan diameter sebesar 2.1 m, efisiensi sebesar 0.421, pitch ratio se-besar 0.591, dengan putaran propeler 400 rpm. Daya mesin yg dibutuhkan barge pada kondisi maksimum (BHPMCR sebesar 1669.5 HP. Dengan mempertimbangkan daya tersebut, maka dipilih mesin jenis Caterpillar tipe Marine 3516B yang mem-punyai daya maksimum sebesar 1285 kilowatt atau 1722.5 horsepower dengan putaran mesin sebesar 1200 rpm

  1. Metal hydride and pyrophoric fuel additives for dicyclopentadiene based hybrid propellants

    Science.gov (United States)

    Shark, Steven C.

    The purpose of this study is to investigate the use of reactive energetic fuel additives that have the potential to increase the combustion performance of hybrid rocket propellants in terms of solid fuel regression rate and combustion efficiency. Additives that can augment the combustion flame zone in a hybrid rocket motor by means of increased energy feedback to the fuel grain surface are of great interest. Metal hydrides have large volumetric hydrogen densities, which gives these materials high performance potential as fuel additives in terms of specifc impulse. The excess hydrogen and corresponding base metal may also cause an increase in the hybrid rocket solid fuel regression rate. Pyrophoric additives also have potential to increase the solid fuel regression rate by reacting more readily near the burning fuel surface providing rapid energy feedback. An experimental performance evaluation of metal hydride fuel additives for hybrid rocket motor propulsion systems is examined in this study. Hypergolic ignition droplet tests and an accelerated aging study revealed the protection capabilities of Dicyclopentadiene (DCPD) as a fuel binder, and the ability for unaided ignition. Static hybrid rocket motor experiments were conducted using DCPD as the fuel. Sodium borohydride (NabH4) and aluminum hydride (AlH3) were examined as fuel additives. Ninety percent rocket grade hydrogen peroxide (RGHP) was used as the oxidizer. In this study, the sensitivity of solid fuel regression rate and characteristic velocity (C*) efficiency to total fuel grain port mass flux and particle loading is examined. These results were compared to HTPB combustion performance as a baseline. Chamber pressure histories revealed steady motor operation in most tests, with reduced ignition delays when using NabH4 as a fuel additive. The addition of NabH4 and AlH3 produced up to a 47% and 85% increase in regression rate over neat DCPD, respectively. For all test conditions examined C* efficiency ranges

  2. Multi-Stage Hybrid Rocket Conceptual Design for Micro-Satellites Launch using Genetic Algorithm

    Science.gov (United States)

    Kitagawa, Yosuke; Kitagawa, Koki; Nakamiya, Masaki; Kanazaki, Masahiro; Shimada, Toru

    The multi-objective genetic algorithm (MOGA) is applied to the multi-disciplinary conceptual design problem for a three-stage launch vehicle (LV) with a hybrid rocket engine (HRE). MOGA is an optimization tool used for multi-objective problems. The parallel coordinate plot (PCP), which is a data mining method, is employed in the post-process in MOGA for design knowledge discovery. A rocket that can deliver observing micro-satellites to the sun-synchronous orbit (SSO) is designed. It consists of an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank and a nozzle. The objective functions considered in this study are to minimize the total mass of the rocket and to maximize the ratio of the payload mass to the total mass. To calculate the thrust and the engine size, the regression rate is estimated based on an empirical model for a paraffin (FT-0070) propellant. Several non-dominated solutions are obtained using MOGA, and design knowledge is discovered for the present hybrid rocket design problem using a PCP analysis. As a result, substantial knowledge on the design of an LV with an HRE is obtained for use in space transportation.

  3. MHD thrust vectoring of a rocket engine

    Science.gov (United States)

    Labaune, Julien; Packan, Denis; Tholin, Fabien; Chemartin, Laurent; Stillace, Thierry; Masson, Frederic

    2016-09-01

    In this work, the possibility to use MagnetoHydroDynamics (MHD) to vectorize the thrust of a solid propellant rocket engine exhaust is investigated. Using a magnetic field for vectoring offers a mass gain and a reusability advantage compared to standard gimbaled, elastomer-joint systems. Analytical and numerical models were used to evaluate the flow deviation with a 1 Tesla magnetic field inside the nozzle. The fluid flow in the resistive MHD approximation is calculated using the KRONOS code from ONERA, coupling the hypersonic CFD platform CEDRE and the electrical code SATURNE from EDF. A critical parameter of these simulations is the electrical conductivity, which was evaluated using a set of equilibrium calculations with 25 species. Two models were used: local thermodynamic equilibrium and frozen flow. In both cases, chlorine captures a large fraction of free electrons, limiting the electrical conductivity to a value inadequate for thrust vectoring applications. However, when using chlorine-free propergols with 1% in mass of alkali, an MHD thrust vectoring of several degrees was obtained.

  4. Enhancement of hybrid rocket combustion performance using nano-sized energetic particles

    Science.gov (United States)

    Risha, Grant Alexander

    Until now, the regression rate of classical hybrid rocket engines have typically been an order of magnitude lower than solid propellant motors; thus, hybrids require a relatively large fuel surface area for a given thrust level. In addition to low linear regression rates, relatively low combustion efficiency (87 to 92%), low mass burning rates, varying oxidizer-to-fuel ratio during operation, and lack of scaling laws have been reported. These disadvantages can be ameliorated by introducing nano-sized energetic powder additives into the solid fuel. The addition of nano-sized energetic particles into the solid fuel enhances performance as measured by parameters such as: density specific impulse, mass and linear burning rates, and thrust. Thermophysical properties of the solid fuel such as density, heat of combustion, thermal diffusivity, and thermal conductivity are also enhanced. The types of nano-sized energetic particles used in this study include aluminum, boron, boron carbide, and some Viton-A coated particles. Since the combustion process of solid fuels in a hybrid rocket engine is governed by the mass flux of the oxidizer entering the combustion chamber, the rate-limiting process is the mixing and reacting of the pyrolysis products of the fuel grain with the incoming oxidizer. The overall goal of this research was to determine the relative propulsive and combustion behavior for a family of newly-developed HTPB-based solid-fuel formulations containing various nano-sized energetic particles. Seventeen formulations contained 13% additive by weight, one formulation (SF4) contained 6.5% additive by weight, and one formulation (SF19) contained 5.65% boron by weight. The two hybrid rocket engines which were used in this investigation were the Long Grain Center-Perforated (LGCP) rocket engine and the X-Ray Transparent Casing (XTC) rocket engine. The smaller scale LGCP rocket engine was used to evaluate all of the formulations because conducting experiments using the

  5. A review of research in low earth orbit propellant collection

    Science.gov (United States)

    Singh, Lake A.; Walker, Mitchell L. R.

    2015-05-01

    This comprehensive review examines the efforts of previous researchers to develop concepts for propellant-collecting spacecraft, estimate the performance of these systems, and understand the physics involved. Rocket propulsion requires the spacecraft to expend two fundamental quantities: energy and propellant mass. A growing number of spacecraft collect the energy they need to execute propulsive maneuvers in-situ with solar panels. In contrast, every spacecraft using rocket propulsion has carried all of the propellant mass needed for the mission from the ground, which limits the range and mission capabilities. Numerous researchers have explored the concept of collecting propellant mass while in space. These concepts have varied in scale and complexity from chemical ramjets to fusion-driven interstellar vessels. Research into propellant-collecting concepts occurred in distinct eras. During the Cold War, concepts tended to be large, complex, and nuclear powered. After the Cold War, concepts transitioned to solar power sources and more effort has been devoted to detailed analysis of specific components of the propellant-collecting architecture. By detailing the major contributions and limitations of previous work, this review concisely presents the state-of-the-art and outlines five areas for continued research. These areas include air-compatible cathode technology, techniques to improve propellant utilization on atmospheric species, in-space compressor and liquefaction technology, improved hypersonic and hyperthermal free molecular flow inlet designs, and improved understanding of how design parameters affect system performance.

  6. Rocket Tablet,

    Science.gov (United States)

    1984-09-12

    the railway platform filled with sand, consumed by a great wave of emotion, looking at the soldiers and waving tearful farewells to them as well as...many matters awaiting me, how can I not be anxious?" Li Fuze rose and said farewell and Li Juemin insisted on -sending him off. These two comrades-in...front of the .? tombstone were the wife of the hero and several hundred solders; an entire rocket troop mourned the hero.... ...... from today on no

  7. Accelerated aging of AP/HTPB propellants and the influence of various environmental aging conditions

    NARCIS (Netherlands)

    Keizers, H.L.J.

    1995-01-01

    Preliminary resuits on accelerated aging of lab-scale produced AP/HTPB propellant and propellants from dissectioned rocket motors are discussed, including aging logic, storage conditions, test techniques and resuits on mechanical, ballistic and safety testing. The mam aging effect observed was

  8. Numerical investigation on the regression rate of hybrid rocket motor with star swirl fuel grain

    Science.gov (United States)

    Zhang, Shuai; Hu, Fan; Zhang, Weihua

    2016-10-01

    Although hybrid rocket motor is prospected to have distinct advantages over liquid and solid rocket motor, low regression rate and insufficient efficiency are two major disadvantages which have prevented it from being commercially viable. In recent years, complex fuel grain configurations are attractive in overcoming the disadvantages with the help of Rapid Prototyping technology. In this work, an attempt has been made to numerically investigate the flow field characteristics and local regression rate distribution inside the hybrid rocket motor with complex star swirl grain. A propellant combination with GOX and HTPB has been chosen. The numerical model is established based on the three dimensional Navier-Stokes equations with turbulence, combustion, and coupled gas/solid phase formulations. The calculated fuel regression rate is compared with the experimental data to validate the accuracy of numerical model. The results indicate that, comparing the star swirl grain with the tube grain under the conditions of the same port area and the same grain length, the burning surface area rises about 200%, the spatially averaged regression rate rises as high as about 60%, and the oxidizer can combust sufficiently due to the big vortex around the axis in the aft-mixing chamber. The combustion efficiency of star swirl grain is better and more stable than that of tube grain.

  9. Burning rate characteristics of energetic CMDB propellants. Part 2. Effect of HMX addition; Ko enerugi CMDB suishin yaku no nensho sokudo tokusei ( II ) - HMX tenka no koka -

    Energy Technology Data Exchange (ETDEWEB)

    Aoki, I. [Nissan Motor Co. Ltd. (Japan)

    1997-08-01

    Burning rate and specific impulse of a solid propellant are extremely important parameters in a design of a solid rocket motor. In this study, the relations between the burning rate and the amount of energy contained in HMX-CMDB propellants wherein the amount of energy is varied by adding HMX (High Melting Point Explosive). The following results are obtained. The final flame temperature is getting higher when the amount of energy is increased by adding HMX into a double-base propellant. The higher the final flame temperature is, the lower the burning rate is. Dark zone temperature, as a physical property, is lowered when the containing amount of energy is increased by adding HMX into the double-base propellant. This is because that, when weight fraction of HMX is increased, reaction heat at burning surface decreases, and the reaction in fizz zone is getting slower. The higher the dark zone temperature is, the higher the burning rate is. 20 refs., 11 figs., 1 tab.

  10. Evaluation of Geopolymer Concrete for Rocket Test Facility Flame Deflectors

    Science.gov (United States)

    Allgood, Daniel C.; Montes, Carlos; Islam, Rashedul; Allouche, Erez

    2014-01-01

    The current paper presents results from a combined research effort by Louisiana Tech University (LTU) and NASA Stennis Space Center (SSC) to develop a new alumina-silicate based cementitious binder capable of acting as a high performance refractory material with low heat ablation rate and high early mechanical strength. Such a binder would represent a significant contribution to NASA's efforts to develop a new generation of refractory 'hot face' liners for liquid or solid rocket plume environments. This project was developed as a continuation of on-going collaborations between LTU and SSC, where test sections of a formulation of high temperature geopolymer binder were cast in the floor and walls of Test Stand E-1 Cell 3, an active rocket engine test stand flame trench. Additionally, geopolymer concrete panels were tested using the NASA-SSC Diagnostic Test Facility (DTF) thruster, where supersonic plume environments were generated on a 1ft wide x 2ft long x 6 inch deep refractory panel. The DTF operates on LOX/GH2 propellants producing a nominal thrust of 1,200 lbf and the combustion chamber conditions are Pc=625psig, O/F=6.0. Data collected included high speed video of plume/panel area and surface profiles (depth) of the test panels measured on a 1-inch by 1-inch giving localized erosion rates during the test. Louisiana Tech conducted a microstructure analysis of the geopolymer binder after the testing program to identify phase changes in the material.

  11. Aircraft Propeller Hub Repair

    Energy Technology Data Exchange (ETDEWEB)

    Muth, Thomas R [ORNL; Peter, William H [ORNL

    2015-02-13

    The team performed a literature review, conducted residual stress measurements, performed failure analysis, and demonstrated a solid state additive manufacturing repair technique on samples removed from a scrapped propeller hub. The team evaluated multiple options for hub repair that included existing metal buildup technologies that the Federal Aviation Administration (FAA) has already embraced, such as cold spray, high velocity oxy-fuel deposition (HVOF), and plasma spray. In addition the team helped Piedmont Propulsion Systems, LLC (PPS) evaluate three potential solutions that could be deployed at different stages in the life cycle of aluminum alloy hubs, in addition to the conventional spray coating method for repair. For new hubs, a machining practice to prevent fretting with the steel drive shaft was recommended. For hubs that were refurbished with some material remaining above the minimal material condition (MMC), a silver interface applied by an electromagnetic pulse additive manufacturing method was recommended. For hubs that were at or below the MMC, a solid state additive manufacturing technique using ultrasonic welding (UW) of thin layers of 7075 aluminum to the hub interface was recommended. A cladding demonstration using the UW technique achieved mechanical bonding of the layers showing promise as a viable repair method.

  12. Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

    OpenAIRE

    Chehroudi, Bruce

    2012-01-01

    Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME) or Vulcain (Ariane 5), the injected li...

  13. Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

    OpenAIRE

    Mboyi, Kalomba; Ren, Junxue; Liu, Yu

    2015-01-01

    A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the e...

  14. Chemical Kinetics Phenomena in Rocket Engines

    Science.gov (United States)

    1952-06-01

    boric acid fluctuation bands are by far the most prominent feature of the spectrum of the exhaust flame from this propellant combination. Weak bands due...spectrum shows only the boric acid fluctuation bands with the super- imposed BO bands of Singh. Some weak bands appear above 6000 A on some of the...accurately. b. Aniline and Furfuryl Alcohol with YFNA. - The spectrum of this rocket flame appears to be continuous in the visible region, though weak bands which

  15. Problems of the mathematical description of rocket engines as plants

    Science.gov (United States)

    Kiforenko, B. N.

    2012-09-01

    Mathematical models of liquid-propellant, nuclear, and electric rocket engines are presented that more fully describe thrust generation than the classical models do. The optimal control of engine thrust is analyzed within the framework of Mayer's general variational problem. It is shown that the control of a rocket engine satisfying the necessary optimality conditions belongs to the boundary arc of the feasible control set between the point of maximum thrust and the point of maximum exhaust velocity

  16. Scientific Experiences Using Argentinean Sounding Rockets in Antarctica

    Science.gov (United States)

    Sánchez-Peña, Miguel

    2000-07-01

    Argentina in the sixties and seventies, had experience for developing and for using sounding rockets and payloads to perform scientific space experiments. Besides they have several bases in Antarctica with adequate premises and installations, also duly equipped aircrafts and trained crews to flight to the white continent. In February 1965, scientists and technical people from the "Instituto de Investigacion Aeronáutica y Espacial" (I.I.A.E.) with the cooperation of the Air Force and the Tucuman University, conducted the "Matienzo Operation" to measure X radiation and temperature in the upper atmosphere, using the Gamma Centauro rocket and also using big balloons. The people involved in the experience, the launcher, other material and equipment flew from the south tip of Argentina to the Matienzo base in Antarctica, in a C-47 aircraft equipped with skies an additional jet engine Marbore 2-C. Other experience was performed in 1975 in the "Marambio" Antartic Base, using the two stages solid propellent sounding rocket Castor, developed in Argentina. The payload was developed in cooperation with the Max Planck Institute of Germany. It consist of a special mixture including a shape charge to form a ionized cloud producing a jet of electrons travelling from Marambio base to the conjugate point in the Northern hemisphere. The cloud was observed by several ground stations in Argentina and also by a NASA aircraft with TV cameras, flying at East of New York. The objective of this experience was to study the electric and magnetic fields in altitude, the neutral points, the temperature and electrons profile. The objectives of both experiments were accomplished satisfactorily.

  17. Liquid-Propellant Droplet Dynamics and Combustion in Supercritical Forced-Convective Environments

    National Research Council Canada - National Science Library

    Yang, Vigor

    1998-01-01

    .... The purpose is to establish a solid theoretical basis for enhancing the understanding of liquid propellant droplet vaporization, combustion, and dynamics at supercritical conditions, with emphasis...

  18. Propellant Preparation Laboratory Complex (Area1-21)

    Data.gov (United States)

    Federal Laboratory Consortium — Description: Area 1-21 is an explosion resistant complex of nine cells built into the side of a granite ridge. Three solid propellant cutting cells are housed in the...

  19. Rocket Engine Altitude Simulation Technologies

    Science.gov (United States)

    Woods, Jody L.; Lansaw, John

    2010-01-01

    John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

  20. Spark Ignition of Combustible Vapor in a Plastic Bottle as a Demonstration of Rocket Propulsion

    Science.gov (United States)

    Mattox, J. R.

    2017-01-01

    I report an innovation that provides a compelling demonstration of rocket propulsion, appropriate for students of physics and other physical sciences. An electrical spark is initiated from a distance to cause the deflagration of a combustible vapor mixed with air in a lightweight plastic bottle that is consequently propelled as a rocket by the…

  1. Mobile propeller dynamometer validation

    Science.gov (United States)

    Morris, Mason Wade

    With growing interest in UAVs and OSU's interest in propeller performance and manufacturing, evaluating UAV propeller and propulsion system performance has become essential. In attempts to evaluate these propellers a mobile propeller dynamometer has been designed, built, and tested. The mobile dyno has been designed to be cost effective through the ability to load it into the back of a test vehicle to create simulated forward flight characteristics. This allows much larger propellers to be dynamically tested without the use of large and expensive wind tunnels. While evaluating the accuracy of the dyno, several improvements had to be made to get accurate results. The decisions made to design and improve the mobile propeller dyno will be discussed along with attempts to validate the dyno by comparing its results against known sources. Another large part of assuring the accuracy of the mobile dyno is determining if the test vehicle will influence the flow going into the propellers being tested. The flow into the propeller needs to be as smooth and uniform as possible. This is determined by characterizing the boundary layer and accelerated flow over the vehicle. This evaluation was accomplished with extensive vehicle aerodynamic measurements with the use of full-scale tests using a pitot-rake and the actual test vehicle. Additional tests were conducted in Oklahoma State University's low speed wind tunnel with a 1/8-scale model using qualitative flow visualization with smoke. Continuing research on the mobile dyno will be discussed, along with other potential uses for the dyno.

  2. Computing Instability In Combustion Of Liquid Propellants

    Science.gov (United States)

    Chen, Yen-Sen; Shang, Huan-Min

    1995-01-01

    Computational fluid dynamics (CFD) code developed for use in design analyses of flow instabilities associated with combustion of sprayed liquid propellants in rocket engines. Code also contributes to design of improved commercial sprayed-fuel combustors in furnaces and jet engines. Proves robust, user-friendly software tool with comprehensive analysis capability. Enables characterization of stability or instability of engine in terms of such physically meaningful parameters as initial conditions of spray, spatial distribution of ratio between concentrations of fuel and oxidizer at injector faces, geometry of combustor, and configurations of baffles.

  3. Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes

    Energy Technology Data Exchange (ETDEWEB)

    Kinefuchi, K. [Department of Aeronautics and Astronautics, University of Tokyo, 7-3-1, Hongo, Bunkyo-ku, Tokyo 113-8656 (Japan); Funaki, I.; Shimada, T.; Abe, T. [Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, 3-1-1, Yoshinodai, Chuo-ku, Sagamihara, Kanagawa 252-5210 (Japan)

    2012-10-15

    Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.

  4. Metallized Gelled Propellant Heat Transfer Tests Analyzed

    Science.gov (United States)

    Palaszewski, Bryan A.

    1997-01-01

    A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted at the NASA Lewis Research Center. These experiments used a small 20- to 40-lbf thrust engine composed of a modular injector, an igniter, a chamber, and a nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt % loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each channel used water flow to carry heat away from the chamber and the attached thermocouples; flow meters allowed heat flux estimates at each of the 31 stations.

  5. Investigation on the foaming behaviors of NC-based gun propellants

    Directory of Open Access Journals (Sweden)

    Yu-xiang Li

    2014-09-01

    Full Text Available To prepare the porous NC-based (nitrocellulose-based gun propellants, the batch foaming process of using supercritical CO2 as the physical blowing agent is used. The solubilities of CO2 in the single-base propellants and TEGDN (trimethyleneglycol dinitrate propellants are measured by the gravimetric method, and SEM (scanning electron microscope is used to observe the morphology of foamed propellants. The result shows that a large amount of CO2 could be dissolved in NC-based propellants. The experimental results also reveal that the energetic plasticizer TEGDN exerts an important influence on the pore structure. The triaxial tensile failure mechanism for solid-state nucleation is used to explain the nucleation of NC-based propellants in the solid state. Since some specific foaming behaviors of NC-based propellants can not be explained by the failure mechanism, a solid-state nucleation mechanism which revises the triaxial tensile failure mechanism is proposed and discussed.

  6. Studies on an aerial propellant transfer space plane (APTSP)

    Energy Technology Data Exchange (ETDEWEB)

    Jayan, N.; Biju Kumar, K.S.; Gupta, A.K.; Kashyap, A.K.; Venkatraman, K.; Mathews, J.; Makunda, H.S. [Indian Institute of Science, Bangalore (India). Dept. of Aerospace Engineering

    2004-03-01

    This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircraft for propellant filling at an altitude of 13 km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10 ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives. (author)

  7. Rocket Scientist for a Day: Investigating Alternatives for Chemical Propulsion

    Science.gov (United States)

    Angelin, Marcus; Rahm, Martin; Gabrielsson, Erik; Gumaelius, Lena

    2012-01-01

    This laboratory experiment introduces rocket science from a chemistry perspective. The focus is set on chemical propulsion, including its environmental impact and future development. By combining lecture-based teaching with practical, theoretical, and computational exercises, the students get to evaluate different propellant alternatives. To…

  8. Analysis of supercritical methane in rocket engine cooling channels

    NARCIS (Netherlands)

    Denies, L.; Zandbergen, B.T.C.; Natale, P.; Ricci, D.; Invigorito, M.

    2016-01-01

    Methane is a promising propellant for liquid rocket engines. As a regenerative coolant, it would be close to its critical point, complicating cooling analysis. This study encompasses the development and validation of a new, open-source computational fluid dynamics (CFD) method for analysis of

  9. Air-Breathing Rocket Engine Test

    Science.gov (United States)

    2000-01-01

    This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  10. Recent Advances and Applications in Cryogenic Propellant Densification Technology

    Science.gov (United States)

    Tomsik, Thomas M.

    2000-01-01

    This purpose of this paper is to review several historical cryogenic test programs that were conducted at the NASA Glenn Research Center (GRC), Cleveland, Ohio over the past fifty years. More recently these technology programs were intended to study new and improved denser forms of liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic rocket fuels. Of particular interest are subcooled cryogenic propellants. This is due to the fact that they have a significantly higher density (eg. triple-point hydrogen, slush etc.), a lower vapor pressure and improved cooling capacity over the normal boiling point cryogen. This paper, which is intended to be a historical technology overview, will trace the past and recent development and testing of small and large-scale propellant densification production systems. Densifier units in the current GRC fuels program, were designed and are capable of processing subcooled LH2 and L02 propellant at the X33 Reusable Launch Vehicle (RLV) scale. One final objective of this technical briefing is to discuss some of the potential benefits and application which propellant densification technology may offer the industrial cryogenics production and end-user community. Density enhancements to cryogenic propellants (LH2, LO2, CH4) in rocket propulsion and aerospace application have provided the opportunity to either increase performance of existing launch vehicles or to reduce the overall size, mass and cost of a new vehicle system.

  11. Some physical and thermodynamic properties of rocket exhaust clouds measured with infrared scanners

    Science.gov (United States)

    Gomberg, R. I.; Kantsios, A. G.; Rosensteel, F. J.

    1977-01-01

    Measurements using infrared scanners were made of the radiation from exhaust clouds from liquid- and solid-propellant rocket boosters. Field measurements from four launches were discussed. These measurements were intended to explore the physical and thermodynamic properties of these exhaust clouds during their formation and subsequent dispersion. Information was obtained concerning the initial cloud's buoyancy, the stabilized cloud's shape and trajectory, the cloud volume as a function of time, and it's initial and stabilized temperatures. Differences in radiation intensities at various wavelengths from ambient and stabilized exhaust clouds were investigated as a method of distinguishing between the two types of clouds. The infrared remote sensing method used can be used at night when visible range cameras are inadequate. Infrared scanning techniques developed in this project can be applied directly to natural clouds, clouds containing certain radionuclides, or clouds of industrial pollution.

  12. Qualification of Magnesium/Teflon/Viton Pyrotechnic Composition Used in Rocket Motors Ignition System

    Directory of Open Access Journals (Sweden)

    Luciana de Barros

    2016-04-01

    Full Text Available The application of fluoropolymers in high-energy-release pyrotechnic compositions is common in the space and defense areas. Pyrotechnic compositions of magnesium/Teflon/Viton are widely used in military flares and pyrogen igniters for igniting the solid propellant of a rocket motor. Pyrotechnic components are considered high-risk products as they may cause catastrophic accidents if initiated or ignited inadvertently. To reduce the hazards involved in the handling, storage and transportation of these devices, the magnesium/Teflon/Viton composition was subjected to various sensitivity tests, DSC and had its stability and compatibility tested with other materials. This composition obtained satisfactory results in all the tests, which qualifies it as safe for production, handling, use, storage and transportation.

  13. Analytical modeling of the gas generator frequency response in hybrid rocket boosters

    Science.gov (United States)

    Osherov, A.; Natan, B.; Gany, A.

    1996-10-01

    Différent types of hybrid engines are considered candidates for space launchers. The gas generator hybrid engine consists of a fuel-rich solid propellant gas generator whose partially burned products are mixed and further burned with a liquid oxidizer in a separate combustion chamber. The objective of this paper is to model and analyze the frequency response characteristics of such a gas generator, as a part of the problem of hybrid rockets longitudinal dynamic instability (POGO). The analysis takes into account the coupling between the gas flow within the grain ports and the elastic body vibrations. In principle, the gas generator admittance coincides with the response of a 'quarter-wave' resonator with dissipative force. It is found that the gas generator structure elasticity affects significantly the GG admittance in both the resonance frequency and the amplification factor.

  14. PARAMETER DESIGN PROPELLER KAPAL

    Directory of Open Access Journals (Sweden)

    Muhammad Ridwan

    2012-04-01

    Full Text Available   Designer propeller kapal harus mempertimbangkan berbagai parameter untuk menghasilkan bentuk, type dan ukuran propeller yang memiliki nilai  efektifitas dan effisiensi propulsi tinggi. Propulsi kapal merupakan faktor yang mendominasi  operasional kapal, karena pemakaian bahan bakar untuk operasional propulsi kapal merupakan 42% dari total cost operasional kapal dan merupakan added value yang akan diperoleh oleh perusahaan pelayaran. Pertimbangan parameter desain propeller argonomis dapat mendukung tujuan di atas sehingga dapat menurunkan pemakaian bahan bakar hingga 20 % saat kapal dioperaionalkan.

  15. Replacement of chemical rocket launchers by beamed energy propulsion.

    Science.gov (United States)

    Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

    2014-11-01

    Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%.

  16. Performance optimization of marine propellers

    Directory of Open Access Journals (Sweden)

    Chang-Sup Lee

    2010-12-01

    In this paper, a design method for increasing performance of the marine propellers including the WCT propeller is suggested. It is described to maximize the performance of the propeller by adjusting expanded areas of the propeller blade. Results show that efficiency can be increased up to over 2% through the suggested design method.

  17. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Science.gov (United States)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

    Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

  18. Hydrodynamics of Ship Propellers

    DEFF Research Database (Denmark)

    Breslin, John P.; Andersen, Poul

    This book deals with flows over propellers operating behind ships, and the hydrodynamic forces and moments which the propeller generates on the shaft and on the ship hull.The first part of the text is devoted to fundamentals of the flow about hydrofoil sections (with and without cavitation......) and about wings. It then treats propellers in uniform flow, first via advanced actuator disc modelling, and then using lifting-line theory. Pragmatic guidance is given for design and evaluation of performance, including the use of computer modelling.The second part covers the development of unsteady forces...... arising from operation in non-uniform hull wakes. First, by a number of simplifications, various aspects of the problem are dealt with separately until the full problem of a non-cavitating, wide-bladed propeller in a wake is treated by a new and completely developed theory. Next, the complicated problem...

  19. Rocket propulsion elements

    CERN Document Server

    Sutton, George P

    2011-01-01

    The definitive text on rocket propulsion-now revised to reflect advancements in the field For sixty years, Sutton's Rocket Propulsion Elements has been regarded as the single most authoritative sourcebook on rocket propulsion technology. As with the previous edition, coauthored with Oscar Biblarz, the Eighth Edition of Rocket Propulsion Elements offers a thorough introduction to basic principles of rocket propulsion for guided missiles, space flight, or satellite flight. It describes the physical mechanisms and designs for various types of rockets' and provides an unders

  20. Modeling of Rocket Fuel Heating and Cooling Processes in the Interior Receptacle Space of Ground-Based Systems

    OpenAIRE

    K. I. Denisova; V. V. Chugunkov

    2016-01-01

    The propellant to fill the fuel tanks of the spacecraft, upper stages, and space rockets on technical and ground-based launch sites before fueling should be prepared to ensure many of its parameters, including temperature, in appropriate condition. Preparation of fuel temperature is arranged through heating and cooling the rocket propellants (RP) in the tanks of fueling equipment. Processes of RP temperature preparation are the most energy-intensive and timeconsuming ones, which require that ...

  1. Manned Mars missions using propellant from space

    Energy Technology Data Exchange (ETDEWEB)

    Zuppero, A.C.; Olson, T.S. (Idaho National Engineering Laboratory, P.O. Box 1625, Idaho Falls, Idaho 83415-3413 (United States)); Redd, L.R. (Department of Energy, Office of Space, Idaho Field Office, Idaho Falls, Idaho 83402 (United States))

    1993-01-10

    .A recent discovery (8/14/92) of a near-earth object containing materials potentially useful for space activities could perhaps change the entire way humans access and operate in space. A near-Earth object ([number sign]4015, 1979 VA, comet Wilson-Harrington) contains water ice that could be used for space propulsion. In addition, this type of object may contain structural and lifesustaining materials (complex hydrocarbons, ammonia and/or bound nitrogen compounds) for space structures, manned planetary bases, or planetary surface terraforming. The retrieval and utilization of rocket propellant from near-Earth objects, for manned Mars missions in particular, has been investigated and the benefits of this scenario to over performing a Mars mission with terrestrial propellants have been documented. The results show water extracted from these objects and retrieved to Earth orbit for use in going to Mars may actually enable manned Mars exploration by reducing the number of Heavy Lift Launch Vehicle (HLLV) flights or eliminating the need for HLLV's altogether. The mission can perhaps be supported with existing launch vehicles and not required heavy lift capability. Also, the development of a nuclear thermal rocket for this alternate approach may be simplified substantially by reducing the operating temperature required.

  2. Rocket Fuel R and D at AFRL: Recent Activities and Future Direction

    Science.gov (United States)

    2017-04-12

    Charts 14 March 2017 - 12 April 2017 Rocket Fuel R&D at AFRL: Recent Activities & Future Direction Matt Billingsley Air Force Research Laboratory (AFMC...Unclassified SAR 31 Matthew Billingsley N/A Q1W8 DISTRIBUTION A: Approved for Public Release; Distribution Unlimited PA Clearance Number 17163 Rocket ... Rocket Kerosene Fuels Chemical Propulsion Liquid Rocket Engines (LRE) Solid Rocket Motors (SRM) Cryogenic (Liquid O2/fuel) Storable (IRFNA, MMH

  3. SRB/SLEEC (Solid Rocket Booster/Shingle Lap Extendible Exit Cone) feasibility study, volume 2. Appendix A: Design study for a SLEEC actuation system

    Science.gov (United States)

    Thompson, D. S.

    1986-01-01

    The results are presented of a design feasibility study of a self-contained (powered) actuation system for a Shingle Lap Extendible Exit Cone (SLEEC) for Transportation System (STS). The evolution of the SLEEC actuation system design is reviewed, the final design concept is summarized, and the results of the detailed study of the final concept of the actuation system are treated. A conservative design using proven mechanical components was established as a major program priority. The final mechanical design has a very low development risk since the components, which consist of ballscrews, gearing, flexible shaft drives, and aircraft cables, have extensive aerospace applications and a history of proven reliability. The mathematical model studies have shown that little or no power is required to deploy the SLEEC actuation system because acceleration forces and internal pressure from the rocket plume provide the required energies. A speed control brake is incorporated in the design in order to control the rate of deployment.

  4. Romanian MRE Rocket Engines Program - An Early Endeavor

    Science.gov (United States)

    Rugescu, R. E.

    2002-01-01

    -channel and 5 digital-channel data measuring and recording system. The first hot test firing of the MRE-1B motor took place successfully on April 9th, 1969 in Bucharest, at the "Elie Carafoli" Chair of UPB. The research program continued with the development of a series of solid, double base propellant rocket and ram-rocket motors, with emphasize on the optimization of the gasdynamic contour of the engine, in order to increase the flight performances. Increments of up to 8% in specific thrust were measured on the test stand, with mass savings and no extra costs. The test firing of the first Romanian, air-breathing ram-rocket engine took place successfully in august 1987 at the Chemical Works in Fagaras, Romania. Astronautics", founded in Bucharest. The principles and history of the "MRE" research program are presented in the proposed paper.

  5. Computational simulation of liquid rocket injector anomalies

    Science.gov (United States)

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.; Davidian, K.

    1986-01-01

    A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

  6. Mechanical Slosh Models for Rocket-Propelled Spacecraft

    Science.gov (United States)

    Jang, Jiann-Woei; Alaniz, Abram; Yang, Lee; Powers. Joseph; Hall, Charles

    2013-01-01

    Several analytical mechanical slosh models for a cylindrical tank with flat bottom are reviewed. Even though spacecrafts use cylinder shaped tanks, most of those tanks usually have elliptical domes. To extend the application of the analytical models for a cylindrical tank with elliptical domes, the modified slosh parameter models are proposed in this report by mapping an elliptical dome cylindrical tank to a flat top/bottom cylindrical tank while maintaining the equivalent liquid volume. For the low Bond number case, the low-g slosh models were also studied. Those low-g models can be used for Bond number > 10. The current low-g slosh models were also modified to extend their applications for the case that liquid height is smaller than the tank radius. All modified slosh models are implemented in MATLAB m-functions and are collected in the developed MST (Mechanical Slosh Toolbox).

  7. Metallic Hydrogen - Potentially a High Energy Rocket Propellant

    Science.gov (United States)

    Cole, John; Silvera, Ike

    2007-01-01

    Pure metallic hydrogen is predicted to have a specific impulse (Isp) of 1700 seconds, but the reaction temperature is too high for current engine materials. Diluting metallic hydrogen with liquid hydrogen can reduce the reaction temperature to levels compatible with current material limits and still provide an Isp greater than 900 s. Metallic hydrogen has not yet been produced on earth, but experimental techniques exist that may change this situation. This paper will provide a brief description of metallic hydrogen and the status of experiments that may soon produce detectable quantities of this material in the lab. Also provided are some characteristics for diluted metallic hydrogen engines and launch vehicles.

  8. The Peak of Rocket Production: The Designer of Ballistic Missiles V.F. Utkin (1923-2000)

    Science.gov (United States)

    Prisniakov, V.; Sitnikova, N.

    2002-01-01

    achievements V. Utkin and his pupils are crea- tion unique "mortar" launching of a heavy liquid rocket from shaft, the decision of a complex of prob- lems on maintenance ready for military action (continuous attendance) of liquid rockets in the filled condi-tion for many years, maintenance of stability of rockets at action on them of striking factors of nuclear explosion. With personal participation of academician IAA V. Utkin the following large scien- tific and technical results were received: (a) a military railway rocket complex with intercontinental solid-propellant rocket with starting weight of 105 tons and with 10 warheads; (b) a method of war manage-ment with the help of command rockets; (c) a method of definition of characteristics of means of overcoming of antimissile defense; (d) war intercontinental rockets with the increased accuracy, with the survivability, with the availability for action; (e) a commanding rocket. Design' decisions not ha- ving the analogues in world: (a) managements of flight solid-propellant an intercontinental ballistic missiles by means of a deviating head part; (b) managements solid-propellant rocket by method of inje- ction of gas in supercritical part of nozzle; (c) industrial introduction of the newest materials etc.V. Ut- kin is the active participant of works in the field of the international cooperation in research and deve- lopment of a space. In 1990 V. Utkin hold a high post of the director of ZSNIIMACH which is leading organization of a space-rocket industry of Russia. Under manual V. Utkin the Federal space program of Russia was developed. V. Utkin had huge authority as the chairman of Advice of the Main designers of the USSR. He was the co-chairman combined commission of experts V. Utkin - T. Stafford" on problems of maintenance joint manned flights. He was the chairman of Coordination advice under the program of researches on manned space complexes. V. Utkin dreamed to be the active participant of a new stage of the outer

  9. Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

    Science.gov (United States)

    Palaszewski, Bryan A.

    1997-01-01

    Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

  10. Current status of free radicals and electronically excited metastable species as high energy propellants

    Science.gov (United States)

    Rosen, G.

    1973-01-01

    A survey is presented of free radicals and electronically excited metastable species as high energy propellants for rocket engines. Nascent or atomic forms of diatomic gases are considered free radicals as well as the highly reactive diatomic triatomic molecules that posess unpaired electrons. Manufacturing and storage problems are described, and a review of current experimental work related to the manufacture of atomic hydrogen propellants is presented.

  11. Models of Non-Stationary Thermodynamic Processes in Rocket Engines Taking into Account a Chemical Equilibrium of Combustion Products

    Directory of Open Access Journals (Sweden)

    A. V. Aliev

    2015-01-01

    Full Text Available The paper considers the two approach-based techniques for calculating the non-stationary intra-chamber processes in solid-propellant rocket engine (SPRE. The first approach assumes that the combustion products are a mechanical mix while the other one supposes it to be the mix, which is in chemical equilibrium. To enhance reliability of solution of the intra ballistic tasks, which assume a chemical equilibrium of combustion products, the computing algorithms to calculate a structure of the combustion products are changed. The algorithm for solving a system of the nonlinear equations of chemical equilibrium, when determining the iterative amendments, uses the orthogonal QR method instead of a method of Gauss. Besides, a possibility to apply genetic algorithms in a task about a structure of combustion products is considered.It is shown that in the tasks concerning the prediction of non-stationary intra ballistic characteristics in a solid propellant rocket engine, application of models of mechanical mix and chemically equilibrium structure of combustion products leads to qualitatively and quantitatively coinciding results. The maximum difference in parameters is 5-10%, at most. In tasks concerning the starting operation of a solid sustainer engine with high-temperature products of combustion difference in results is more essential, and can reach 20% and more.A technique to calculate the intra ballistic parameters, in which flotation of combustion products is considered in the light of a spatial statement, requires using the high-performance computer facilities. For these tasks it is offered to define structure of products of combustion and its thermo-physical characteristics, using the polynoms coefficients of which should be predefined.

  12. Advanced propeller research

    Science.gov (United States)

    Groeneweg, John F.; Bober, Lawrence J.

    1990-01-01

    Recent results of aerodynamic and acoustic research on both single rotation and counterrotation propellers are reviewed. Data and analytical results are presented for three propellers: SR-7A, the single rotation design used in the NASA Propfan Test Assessment (PTA) flight program; CRP-X1, the initial 5+5 Hamilton Standard counterrotating design; and F7-A7, the 8+8 counterrotating G.E. design used in the proof of concept Unducted Fan (UDF) engine. In addition to propeller efficiencies, cruise and takeoff noise, and blade pressure data, off-design phenomena involving formation of leading edge vortexes are described. Aerodynamic and acoustic computational results derived from 3-D Euler and acoustic radiation codes are presented. Research on unsteady flows which are particularly important for understanding counterrotation interaction noise, unsteady loading effects on acoustics, and flutter or forced response is described. The first results of 3-D unsteady Euler solutions are illustrated for a single rotation propeller at angle of attack and for a counterrotation propeller. Basic experimental and theoretical results from studies on the unsteady aerodynamics of oscillating cascades are outlined.

  13. Rockets two classic papers

    CERN Document Server

    Goddard, Robert

    2002-01-01

    Rockets, in the primitive form of fireworks, have existed since the Chinese invented them around the thirteenth century. But it was the work of American Robert Hutchings Goddard (1882-1945) and his development of liquid-fueled rockets that first produced a controlled rocket flight. Fascinated by rocketry since boyhood, Goddard designed, built, and launched the world's first liquid-fueled rocket in 1926. Ridiculed by the press for suggesting that rockets could be flown to the moon, he continued his experiments, supported partly by the Smithsonian Institution and defended by Charles Lindbergh. T

  14. On fundamentally new sources of energy for rockets in the early works of the pioneers of astronautics

    Science.gov (United States)

    Melkumov, T. M.

    1977-01-01

    The research for more efficient methods of propelling a spacecraft, than can be achieved with chemical energy, was studied. During a time when rockets for space flight had not actually been built pioneers in rocket technology were already concerned with this problem. Alternative sources proposed at that time, were nuclear and solar energy. Basic engineering problems of each source were investigated.

  15. Propeller TAP flap

    DEFF Research Database (Denmark)

    Thomsen, Jørn Bo; Bille, Camilla; Wamberg, Peter

    2013-01-01

    The aim of this study was to examine if a propeller thoracodorsal artery perforator (TAP) flap can be used for breast reconstruction. Fifteen women were reconstructed using a propeller TAP flap, an implant, and an ADM. Preoperative colour Doppler ultrasonography was used for patient selection...... to identify the dominant perforator in all cases. A total of 16 TAP flaps were performed; 12 flaps were based on one perforator and four were based on two. A permanent silicone implant was used in 14 cases and an expander implant in two. Minor complications were registered in three patients. Two cases had...... major complications needing additional surgery. One flap was lost due to a vascular problem. Breast reconstruction can be performed by a propeller TAP flap without cutting the descending branch of the thoracodorsal vessels. However, the authors would recommend that a small cuff of muscle is left around...

  16. Laser Ignition Characteristics of GOX/GH2 and GOX/GCH4 Propellants

    OpenAIRE

    長谷川, 恵一; HASEGAWA, Keichi; 佐藤, 政裕; Sato, Masahiro

    2003-01-01

    The laser ignition characteristics of gaseous oxygen (GOX) / gaseous hydrogen (GH2) and GOX / gaseous methane (GCH4) as non-toxic propellants have been under investigation at the Rocket Propulsion Center. Because of the advantages of initiating ignition with laser beam transmitted through an optical fiber, this system is expected to be used in liquid rocket engines with multiple combustion chambers, such as the RCS (Reaction Control System) and the aero-spike nozzle engine. Another advantage ...

  17. The sky is falling II: Impact of deposition produced during the static testing of solid rocket motors on corn and alfalfa.

    Science.gov (United States)

    Doucette, William J; Mendenhall, Scout; McNeill, Laurie S; Heavilin, Justin

    2014-06-01

    Tests of horizontally restrained rocket motors at the ATK facility in Promontory, Utah, USA result in the deposition of an estimated 1.5million kg of entrained soil and combustion products (mainly aluminum oxide, gaseous hydrogen chloride and water) on the surrounding area. The deposition is referred to as test fire soil (TFS). Farmers observing TFS deposited on their crops expressed concerns regarding the impact of this material. To address these concerns, we exposed corn and alfalfa to TFS collected during a September 2009 test. The impact was evaluated by comparing the growth and tissue composition of controls relative to the treatments. Exposure to TFS, containing elevated levels of chloride (1000 times) and aluminum (2 times) relative to native soils, affected the germination, growth and tissue concentrations of various elements, depending on the type and level of exposure. Germination was inhibited by high concentrations of TFS in soil, but the impact was reduced if the TFS was pre-leached with water. Biomass production was reduced in the TFS amended soils and corn grown in TFS amended soils did not develop kernels. Chloride concentrations in corn and alfalfa grown in TFS amended soils were two orders of magnitude greater than controls. TFS exposed plants contained higher concentrations of several cations, although the concentrations were well below livestock feed recommendations. Foliar applications of TFS had no impact on biomass, but some differences in the elemental composition of leaves relative to controls were observed. Washing the TFS off the leaves lessened the impact. Results indicate that the TFS deposition could have an effect, depending on the amount and growth stage of the crops, but the impact could be mitigated with rainfall or the application of additional irrigation water. The high level of chloride associated with the TFS is the main cause of the observed impacts. Copyright © 2014 Elsevier B.V. All rights reserved.

  18. Refinement of Propellant Strand Burning Method to Suit Aluminised Composite Rocket Propellant

    Science.gov (United States)

    2014-12-01

    with a soft brush (dedusting) and the strands were inhibited with paint diluted by water. The paint was a Super Flat acrylic, deep tint base (Line 500...Unfortunately, the viscosity of the epoxy was such that the excess needed to be manually removed by scraping the strands on the edges of the bath as they were...ANNOUNCEMENT No Limitations 17. CITATION IN OTHER DOCUMENTS Yes 18. DSTO RESEARCH LIBRARY THESAURUS http:// web -vic.dsto.defence.gov.au/workareas

  19. Propeller torque load and propeller shaft torque response correlation during ice-propeller interaction

    Science.gov (United States)

    Polić, Dražen; Ehlers, Sören; Æsøy, Vilmar

    2017-03-01

    Ships use propulsion machinery systems to create directional thrust. Sailing in ice-covered waters involves the breaking of ice pieces and their submergence as the ship hull advances. Sometimes, submerged ice pieces interact with the propeller and cause irregular fluctuations of the torque load. As a result, the propeller and engine dynamics become imbalanced, and energy propagates through the propulsion machinery system until equilibrium is reached. In such imbalanced situations, the measured propeller shaft torque response is not equal to the propeller torque. Therefore, in this work, the overall system response is simulated under the ice-related torque load using the Bond graph model. The energy difference between the propeller and propeller shaft is estimated and related to their corresponding mechanical energy. Additionally, the mechanical energy is distributed among modes. Based on the distribution, kinetic and potential energy are important for the correlation between propeller torque and propeller shaft response.

  20. Erosive Burning of Composite Solid Propellants

    Science.gov (United States)

    1979-02-01

    motion picture camera through a series of four quartz windows located along the length of the test section. Frame by frame analysis of the films permits...increased heat transfer from a "core" gas flow (not- 5 ably the widely used model of L. noir and Robillard( 1 4 )). If this is indeed the controlling

  1. The Dynamics of Solid Propellant Combustion.

    Science.gov (United States)

    1980-12-31

    prcoposedl mechanism for os;cillaitory burning of corda Le is the successivc explo- sion, assisted by gas phase heat feedback, of diFsCrete_ sur- face layers...13 1’ 3’’ ]-y 1.3 tL.’’ -’ 4 ,c u onc t.3 (5>2 Cn vo ry n0’rn o101: 1FA-oIt c .v tIIr01 tCi, d ’n e 4𔃽 r3 ( -I_’ j -)I ’.- if󈧥 1 1 (3 1hnl a -’f c

  2. Nonlinear Burning Stability of Solid Propellants.

    Science.gov (United States)

    1983-12-31

    Carlevaro R., and Poli A., "Studio Speri- mentale delle Caratteristiche di Combustione di un Pro pellente Solido a Doppia Base in Regime Non Staziona...34VelocitA Stazionaria di Combustione di un Propellente Solido Alluminiato in funzione di Pres sione e Temperatura", Master’s Thesis in Aeronautical

  3. Hydrocarbon polymeric binder for advanced solid propellant

    Science.gov (United States)

    Potts, J. E. (Editor)

    1972-01-01

    A series of DEAB initiated isoprene polymerizations were run in the 5-gallon stirred autoclave reactor. Polymerization run parameters such as initiator concentration and feed rate were correlated with the molecular weight to provide a basis for molecular weight control in future runs. Synthetic methods were developed for the preparation of n-1,3-alkadienes. By these methods, 1,3-nonadiene was polymerized using DEAB initiator to give an ester-telechelic polynonadiene. This was subsequently hydrogenated with copper chromite catalyst to give a hydroxyl terminated saturated liquid hydrocarbon prepolymer having greatly improved viscosity characteristics and a Tg 18 degrees lower than that of the hydrogenated polyisoprenes. The hydroxyl-telechelic saturated polymers prepared by the hydrogenolysis of ester-telechelic polyisoprene were reached with diisocyanates under conditions favoring linear chain extension gel permeation chromatography was used to monitor this condensation polymerization. Fractions having molecular weights above one million were produced.

  4. Dynamic Regime of Ignition of Solid Propellant

    Directory of Open Access Journals (Sweden)

    Zolotorev Nikolay

    2016-01-01

    Full Text Available This article presents a dynamic regime of exposure of the radiant flux on the sample of gun-cotton. Obtained time the ignition of gun-cotton in the heating conditions of increasing heat flux in the range from 0.2 W/cm2 to 22 W/cm2. A comparison of the delay times of the ignition when heated variable and constant heat flux.

  5. Laser Ignition of Solid Propellant Formulations

    Science.gov (United States)

    1979-06-01

    experiments of Capellos with nitromine explosives have shown that in many cases, confinement of the flame by means of glass is required in order to obtain...I. C. Capellos , US ArM Armament tesearch and Development Coamand. Dover. NJ.private cmmunication. GEO-CENTERS. INC

  6. Hull-Propeller Interaction and Its Effect on Propeller Cavitation

    DEFF Research Database (Denmark)

    Regener, Pelle Bo

    , which presumably does not reflect the differences between the propellers sufficiently. Obtaining effective wake fields using the hybrid RANS-BEM approach at model and full scale also provides the opportunity to investigate the behind-ship cavitation performance of propellers with comparably low...... computational effort. The boundary element method for propeller analysis includes a partially nonlinear cavitation model, which is able to predict partial sheet cavitation and supercavitation. The cavitation behaviour of the conventional propeller and the Kappel propeller from the earlier simulations...... a major effect on propeller cavitation, signifying the importance of using the correct inflow, i.e. the effective wake field when evaluating propeller cavitation performance....

  7. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

    Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL... Rocket Lab Distribution A: Approved for Public Release; Distribution Unlimited. PA Clearance 16493 2Distribution A: Approved for Public Release

  8. The flight of uncontrolled rockets

    CERN Document Server

    Gantmakher, F R; Dryden, H L

    1964-01-01

    International Series of Monographs on Aeronautics and Astronautics, Division VII, Volume 5: The Flight of Uncontrolled Rockets focuses on external ballistics of uncontrolled rockets. The book first discusses the equations of motion of rockets. The rocket as a system of changing composition; application of solidification principle to rockets; rotational motion of rockets; and equations of motion of the center of mass of rockets are described. The text looks at the calculation of trajectory of rockets and the fundamentals of rocket dispersion. The selection further focuses on the dispersion of f

  9. Parametric study and performance analysis of hybrid rocket motors with double-tube configuration

    Science.gov (United States)

    Yu, Nanjia; Zhao, Bo; Lorente, Arnau Pons; Wang, Jue

    2017-03-01

    The practical implementation of hybrid rocket motors has historically been hampered by the slow regression rate of the solid fuel. In recent years, the research on advanced injector designs has achieved notable results in the enhancement of the regression rate and combustion efficiency of hybrid rockets. Following this path, this work studies a new configuration called double-tube characterized by injecting the gaseous oxidizer through a head end injector and an inner tube with injector holes distributed along the motor longitudinal axis. This design has demonstrated a significant potential for improving the performance of hybrid rockets by means of a better mixing of the species achieved through a customized injection of the oxidizer. Indeed, the CFD analysis of the double-tube configuration has revealed that this design may increase the regression rate over 50% with respect to the same motor with a conventional axial showerhead injector. However, in order to fully exploit the advantages of the double-tube concept, it is necessary to acquire a deeper understanding of the influence of the different design parameters in the overall performance. In this way, a parametric study is carried out taking into account the variation of the oxidizer mass flux rate, the ratio of oxidizer mass flow rate injected through the inner tube to the total oxidizer mass flow rate, and injection angle. The data for the analysis have been gathered from a large series of three-dimensional numerical simulations that considered the changes in the design parameters. The propellant combination adopted consists of gaseous oxygen as oxidizer and high-density polyethylene as solid fuel. Furthermore, the numerical model comprises Navier-Stokes equations, k-ε turbulence model, eddy-dissipation combustion model and solid-fuel pyrolysis, which is computed through user-defined functions. This numerical model was previously validated by analyzing the computational and experimental results obtained for

  10. Method and apparatus to produce high specific impulse and moderate thrust from a fusion-powered rocket engine

    Science.gov (United States)

    Cohen, Samuel A.; Pajer, Gary A.; Paluszek, Michael A.; Razin, Yosef S.

    2017-11-21

    A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system for heating a stable plasma to produce fusion reactions in the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. The propellant injection system injects cold propellant into a gas box at one end of the reactor chamber, where the propellant is ionized into a plasma. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust.

  11. Use of Several Thermal Analysis Techniques on a Hypalon Paint Coating for the Solid Rocket Booster (SRB) of the Space Shuttle

    Science.gov (United States)

    Wingard, Charles D.; Whitaker, Ann F. (Technical Monitor)

    2000-01-01

    White Hypalon paint is brush-applied as a moisture barrier coating over cork surfaces on each of the two Space Shuttle SRBs. Fine cracks have been observed in the Hypalon coating three times historically on laboratory witness panels, but never on flight hardware. Samples of the cracked and standard ("good") Hypalon were removed from witness panel cork surfaces, and were tested in 1998 by Thermogravimetric Analysis (TGA), TMA and Differential Scanning Calorimetry (DSC) thermal analysis techniques. The TGA data showed that at 700C, where only paint pigment solids remain, the cracked material had about 9 weight percent more material remaining than the standard material, probably indicating incomplete mixing of the paint before it was brush-applied to produce the cracked material. Use of the TMA film/fiber technique showed that the average modulus (stiffness) vs. temperature was about 3 to 6 times higher for the cracked material than for the standard material. The TMA data also showed that an increase in coating thickness for the cracked Hypalon was not a factor in the anomaly.

  12. Optimum Disposition of Metal Particles in the Propellant Grain

    Directory of Open Access Journals (Sweden)

    Leonid L. Minkov

    2014-01-01

    Full Text Available Using the dispersed metal in solid propellants to increase the temperature of combustion products leads to such a problem as the specific impulse loss due to the incomplete combustion of metal particles in the exhaust products. A redistribution of metal loaded into the propellant grain is one of the methods to decrease the specific impulse loss. This paper reports on the ways to obtain the optimum metal particle disposition for the case-bounded propellant grain of tube cross-sectional type. Three different approaches to analyze the metal combustion efficiency are discussed. The influence of the dynamic nonequilibrium of two-phase flow on the optimum metal particles disposition in the propellant grain of tube cross-sectional type is investigated.

  13. Influence of heat transfer on high pressure flame structure and stabilization in liquid rocket engines

    OpenAIRE

    Mari, Raphaël

    2015-01-01

    This research work deals with the problem of the flame stabilization in the context of high pressure liquid rocket engines. Flame stabilization in a rocket engine is a critical feature. An instability can lead to important damages of the engine or the destruction of the launcher and the satellite. The engines (Vulcain 2 and Vinci) of the Ariane 5, and the future Ariane 6, use the hydrogen/oxygen propellants. One characteristic of this couple is its high specific impulse. The launcher performa...

  14. The liquid rocket booster as an element of the U.S. national space transportation system

    Science.gov (United States)

    Bialla, Paul H.; Simon, Michael C.

    Liquid rocket boosters (LRBs) were first considered for the U.S. Space Transportation System (STS) during the early conceptual phases of the Space Shuttle program. However, solid rocket boosters (SRBs) were ultimately selected for the STS, primarily due to near-term economics. Liquid rocket boosters are once again being considered as a possible future upgrade to the Shuttle. This paper addresses the findings of these studies to date, with emphasis on the feasibility, benefits, and implementation strategy for a LRB program. The principal issue relating to LRB feasibility is their ability to be integrated into the STS with minimal vehicle and facility impacts. Booster size has been shown to have a significant influence on compatibility with the STS. The physical dimensions of the Orbiter and STS support facilities place an inherent limitation on the size of any booster to be used with this system. In addition, excessively large diameter boosters can cause increased airloads to be induced on the Orbiter wings, requiring modification of STS launch trajectory and possible performance losses. However, trajectory and performance analyses have indicated that LRBs can be designed within these sizing constraints and still have sufficient performance to meet Space Shuttle mission requirements. In fact, several configurations have been developed to meet a design goal of providing a 20,000 lb performance improvement to low Earth-orbit (LEO), as compared with current SRBs. Several major system trade studies have been performed to establish a baseline design which is most compatible with the existing Space Transportation System. These trades include propellant selection (storable, hydrogen-oxygen, hydrocarbon-oxygen, and advanced propellants); pump-fed vs pressure-fed propellant feed system design; engine selection (Space Shuttle Main Engine, Titan LR-87, and advanced new engines); number of engines per booster; and reusability vs expendability. In general, it was determined

  15. HIGH ENERGY REPLACEMENT FOR TEFLON PROPELLANT IN PULSED PLASMA THRUSTERS Project

    Data.gov (United States)

    National Aeronautics and Space Administration — This program will utilize a well-characterized Pulsed Plasma Thruster (PPT) to test experimental high-energy extinguishable solid propellants (HE), instead of...

  16. Simulating the Composite Propellant Manufacturing Process

    Science.gov (United States)

    Williamson, Suzanne; Love, Gregory

    2000-01-01

    There is a strategic interest in understanding how the propellant manufacturing process contributes to military capabilities outside the United States. The paper will discuss how system dynamics (SD) has been applied to rapidly assess the capabilities and vulnerabilities of a specific composite propellant production complex. These facilities produce a commonly used solid propellant with military applications. The authors will explain how an SD model can be configured to match a specific production facility followed by a series of scenarios designed to analyze operational vulnerabilities. By using the simulation model to rapidly analyze operational risks, the analyst gains a better understanding of production complexities. There are several benefits of developing SD models to simulate chemical production. SD is an effective tool for characterizing complex problems, especially the production process where the cascading effect of outages quickly taxes common understanding. By programming expert knowledge into an SD application, these tools are transformed into a knowledge management resource that facilitates rapid learning without requiring years of experience in production operations. It also permits the analyst to rapidly respond to crisis situations and other time-sensitive missions. Most importantly, the quantitative understanding gained from applying the SD model lends itself to strategic analysis and planning.

  17. Another Look at Rocket Thrust

    Science.gov (United States)

    Hester, Brooke; Burris, Jennifer

    2012-01-01

    Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force--the thrust--on the rocket. Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of…

  18. Modeling and Fault Simulation of Propellant Filling System

    Science.gov (United States)

    Jiang, Yunchun; Liu, Weidong; Hou, Xiaobo

    2012-05-01

    Propellant filling system is one of the key ground plants in launching site of rocket that use liquid propellant. There is an urgent demand for ensuring and improving its reliability and safety, and there is no doubt that Failure Mode Effect Analysis (FMEA) is a good approach to meet it. Driven by the request to get more fault information for FMEA, and because of the high expense of propellant filling, in this paper, the working process of the propellant filling system in fault condition was studied by simulating based on AMESim. Firstly, based on analyzing its structure and function, the filling system was modular decomposed, and the mathematic models of every module were given, based on which the whole filling system was modeled in AMESim. Secondly, a general method of fault injecting into dynamic system was proposed, and as an example, two typical faults - leakage and blockage - were injected into the model of filling system, based on which one can get two fault models in AMESim. After that, fault simulation was processed and the dynamic characteristics of several key parameters were analyzed under fault conditions. The results show that the model can simulate effectively the two faults, and can be used to provide guidance for the filling system maintain and amelioration.

  19. The gravitational wave rocket

    OpenAIRE

    Bonnor, W. B.; Piper, M. S.

    1997-01-01

    Einstein's equations admit solutions corresponding to photon rockets. In these a massive particle recoils because of the anisotropic emission of photons. In this paper we ask whether rocket motion can be powered only by the emission of gravitational waves. We use the double series approximation method and show that this is possible. A loss of mass and gain in momentum arise in the second approximation because of the emission of quadrupole and octupole waves.

  20. Atomic hydrogen as a launch vehicle propellant

    Science.gov (United States)

    Palaszewski, Bryan A.

    1990-01-01

    An analysis of several atomic hydrogen launch vehicles was conducted. A discussion of the facilities and the technologies that would be needed for these vehicles is also presented. The Gross Liftoff Weights (GLOW) for two systems were estimated; their specific impulses (I sub sp) were 750 and 1500 lb(sub f)/s/lb(sub m). The atomic hydrogen launch vehicles were also compared to the currently planned Advanced Launch System design concepts. Very significant GLOW reductions of 52 to 58 percent are possible over the Advanced Launch System designs. Applying atomic hydrogen propellants to upper stages was also considered. Very high I(sub sp) (greater than 750 lb(sub f)/s/lb(sub m)) is needed to enable a mass savings over advanced oxygen/hydrogen propulsion. Associated with the potential benefits of high I(sub sp) atomic hydrogen are several challenging problems. Very high magnetic fields are required to maintain the atomic hydrogen in a solid hydrogen matrix. The magnetic field strength was estimated to be 30 kilogauss (3 Tesla). Also the storage temperature of the propellant is 4 K. This very low temperature will require a large refrigeration facility for the launch vehicle. The design considerations for a very high recombination rate for the propellant are also discussed. A recombination rate of 210 cm/s is predicted for atomic hydrogen. This high recombination rate can produce very high acceleration for the launch vehicle. Unique insulation or segmentation to inhibit the propellant may be needed to reduce its recombination rate.

  1. Combustion of Solid Propellants (La Combustion des Propergols Solides)

    Science.gov (United States)

    1991-07-01

    poudres modificateurs balistiques, des noirs de A mouler de caract~ristiques volontairement carbone sont souvent incorpor~s car ils ont diffdrontes ce...de antilueurs (suppresseurs de la post noir de carbone eat souvent utilisd combustion). Ceux-ci, g~ndralement des sels industriellemont comae param...devient compressible at is vitesse de combustion Pour 6tablir un bilan sur le risque eat modifide. La figure 7 illustre cet d’appariticn d’instabilit~s

  2. High Frequency Combustion Instabilities of LOx/CH4 Spray Flames in Rocket Engine Combustion Chambers

    NARCIS (Netherlands)

    Sliphorst, M.

    2011-01-01

    Ever since the early stages of space transportation in the 1940’s, and the related liquid propellant rocket engine development, combustion instability has been a major issue. High frequency combustion instability (HFCI) is the interaction between combustion and the acoustic field in the combustion

  3. Space Shuttle Solid Rocket Booster Debris Assessment

    Science.gov (United States)

    Kendall, Kristin; Kanner, Howard; Yu, Weiping

    2006-01-01

    The Space Shuttle Columbia Accident revealed a fundamental problem of the Space Shuttle Program regarding debris. Prior to the tragedy, the Space Shuttle requirement stated that no debris should be liberated that would jeopardize the flight crew and/or mission success. When the accident investigation determined that a large piece of foam debris was the primary cause of the loss of the shuttle and crew, it became apparent that the risk and scope of - damage that could be caused by certain types of debris, especially - ice and foam, were not fully understood. There was no clear understanding of the materials that could become debris, the path the debris might take during flight, the structures the debris might impact or the damage the impact might cause. In addition to supporting the primary NASA and USA goal of returning the Space Shuttle to flight by understanding the SRB debris environment and capability to withstand that environment, the SRB debris assessment project was divided into four primary tasks that were required to be completed to support the RTF goal. These tasks were (1) debris environment definition, (2) impact testing, (3) model correlation and (4) hardware evaluation. Additionally, the project aligned with USA's corporate goals of safety, customer satisfaction, professional development and fiscal accountability.

  4. Solid rocket booster joint seal analyses

    Science.gov (United States)

    Moore, Carleton J.

    1987-01-01

    O-ring response and sealing in pressurized shell structures is examined. The study found that the key elements in the failure of the seal to be joint opening and rotation, assembly out of roundness, and O-ring seal response.

  5. Cryogenic Propellant Storage and Transfer Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Cryogenic Propellant Storage and Transfer project will demonstrate the capability to safely and efficiently store, transfer and measure cryogenic propellants,...

  6. The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

    Science.gov (United States)

    Zubrin, Robert M.

    1991-01-01

    The following paper reports on a design study of a novel space transportation concept known as a 'NIMF' (Nuclear rocket using Indigenous Martian Fuel). The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO2. This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34,000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface.

  7. Composite propellant aluminum agglomeration reduction using tailored Al/PTFE particles

    Science.gov (United States)

    Sippel, Travis R.

    can be ignited via optical flash. Propellant aluminum agglomeration is assessed through replacement of reference aluminum powders (spherical, flake, or nanoscale) with Al/PTFE (90/10 or 70/30 wt.%) particles. The effects on burning rate, pressure dependence, and aluminum ignition, combustion, and agglomeration are quantified. Microscopic imaging shows tailored particles promptly ignite at the burning surface and appear to breakup into smaller particles. Replacement of spherical aluminum with Al/PTFE 70/30 wt.% also increases the pressure exponent from 0.36 to 0.58, which results in a 50% increase in propellant burning rate at 13.8 MPa. Combustion products were quench collected using a liquid-free technique at 2.1 and 6.9 MPa. Sizing of products indicates that composite particles result in nominally 25 μm coarse products, which are smaller than the original, average particle size and are also 66% smaller in diameter (96% by volume) than the 76 μm products collected from reference spherical aluminized propellant. Smaller diameter condensed phase products and more gaseous products will likely decrease two-phase flow loss and reduce slag accumulation in solid rocket motors.

  8. US Rocket Propulsion Industrial Base Health Metrics

    Science.gov (United States)

    Doreswamy, Rajiv

    2013-01-01

    The number of active liquid rocket engine and solid rocket motor development programs has severely declined since the "space race" of the 1950s and 1960s center dot This downward trend has been exacerbated by the retirement of the Space Shuttle, transition from the Constellation Program to the Space launch System (SLS) and similar activity in DoD programs center dot In addition with consolidation in the industry, the rocket propulsion industrial base is under stress. To Improve the "health" of the RPIB, we need to understand - The current condition of the RPIB - How this compares to past history - The trend of RPIB health center dot This drives the need for a concise set of "metrics" - Analogous to the basic data a physician uses to determine the state of health of his patients - Easy to measure and collect - The trend is often more useful than the actual data point - Can be used to focus on problem areas and develop preventative measures The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs. center dot The RPIB encompasses US government, academic, and commercial (including industry primes and their supplier base) research, development, test, evaluation, and manufacturing capabilities and facilities. center dot The RPIB includes the skilled workforce, related intellectual property, engineering and support services, and supply chain operations and management. This definition touches the five main segments of the U.S. RPIB as categorized by the USG: defense, intelligence community, civil government, academia, and commercial sector. The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs

  9. Ignition of Composite Propellants under Rapid Pressure Loading

    Science.gov (United States)

    1981-09-01

    the entrance of the crack cavity, and in the main chamber. Piezoelectric quartz transducers with a rise time of 1.5 us and natural frequency of 300...0C), which is the accepted decomposition temperature of AP.14 3 Because of the heterogeneous nature of the propellant, the AP and the fuel binder...Combustion in Narrow Cavities. J. of Appl. Mech., Vol. 11, 1970, pp. 276-281. 5. Godai, T.: Flame Propagacion into the Crack of a Solid-Propellant Grain

  10. Producing propellants from water in lunar soil using solar lasers

    Science.gov (United States)

    de Morais Mendonca Teles, Antonio

    The exploration of the Solar System is directly related to the efficiency of engines designed to explore it, and consequently, to the propulsion techniques, materials and propellants for those engines. With the present day propulsion techniques it is necessary great quantities of propellants to impulse a manned spacecraft to Mars and beyond in the Solar System, which makes these operations financially very expensive because of the costs involved in launching it from planet Earth, due to its high gravity field strength. To solve this problem, it is needed a planetary place with smaller gravity field strength, near to the Earth and with great quantities of substances at the surface necessary for the in-situ production of propellants for spacecrafts. The only place available is Earth's natural satellite the Moon. So, here in this paper, I propose the creation of a Lunar Propellant Manufacturer. It is a robot-spacecraft which can be launched from Earth using an Energia Rocket, and to land on the Moon in an area (principally near to the north pole where it was discovered water molecules ice recently) with great quantities of oxygen and hydrogen (propellants) in the silicate soil, previously observed and mapped by spacecrafts in lunar orbit, for the extraction of those molecules from the soil and the in-situ production of the necessary propellants. The Lunar Propellant Manufacturer (LPM) spacecraft consists of: 1) a landing system with four legs (extendable) and rovers -when the spacecraft touches down, the legs retract in order that two apparatuses, analogue to tractor's wheeled belts parallel sided and below the spacecraft, can touch firmly the ground -it will be necessary for the displacement of the spacecraft to new areas with richer propellants content, when the early place has already exhausted in propellants; 2) a digging machine -a long, resistant extendable arm with an excavator hand, in the outer part of the spacecraft -it will extend itself to the ground

  11. ROCKET PORT CLOSURE

    Science.gov (United States)

    Mattingly, J.T.

    1963-02-12

    This invention provides a simple pressure-actuated closure whereby windowless observation ports are opened to the atmosphere at preselected altitudes. The closure comprises a disk which seals a windowless observation port in rocket hull. An evacuated instrument compartment is affixed to the rocket hull adjacent the inner surface of the disk, while the outer disk surface is exposed to the atmosphere through which the rocket is traveling. The pressure differential between the evacuated instrument compartment and the relatively high pressure external atmosphere forces the disk against the edge of the observation port, thereby effecting a tight seai. The instrument compartment is evacuated to a pressure equal to the atmospheric pressure existing at the altitude at which it is desiretl that the closure should open. When the rocket reaches this preselected altitude, the inwardly directed atmospheric force on the disk is just equaled by the residual air pressure force within the instrument compartment. Consequently, the closure disk falls away and uncovers the open observation port. The separation of the disk from the rocket hull actuates a switch which energizes the mechanism of a detecting instrument disposed within the instrument compartment. (AE C)

  12. Liquid-propellant droplet vaporization and combustion in high pressure environments

    Science.gov (United States)

    Yang, Vigor

    1991-01-01

    In order to correct the deficiencies of existing models for high-pressure droplet vaporization and combustion, a fundamental investigation into this matter is essential. The objective of this research are: (1) to acquire basic understanding of physical and chemical mechanisms involved in the vaporization and combustion of isolated liquid-propellant droplets in both stagnant and forced-convective environments; (2) to establish droplet vaporization and combustion correlations for the study of liquid-propellant spray combustion and two-phase flowfields in rocket motors; and (3) to investigate the dynamic responses of multicomponent droplet vaporization and combustion to ambient flow oscillations.

  13. Origin of how steam rockets can reduce space transport cost by orders of magnitude

    Science.gov (United States)

    Zuppero, Anthony; Larson, Thomas K.; Schnitzler, Bruce G.; Werner, James E.; Rice, John W.; Hill, Thomas J.; Richins, William D.; Parlier, Lynn

    1999-01-01

    A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10% of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon.

  14. Origin of how steam rockets can reduce space transport cost by orders of magnitude

    Energy Technology Data Exchange (ETDEWEB)

    Zuppero, A.; Larson, T.K.; Schnitzler, B.G.; Rice, J.W.; Hill, T.J.; Richins, W.D.; Parlier, L. [Lockheed Martin Idaho Technologies, Idaho National Engineering and Environmental Laboratory, Idaho Falls, Idaho 83415 (United States); Werner, J.E. [Department of Energy, Idaho Field Office, Idaho National Engineering and Environmental Laboratory, Idaho Falls, Idaho 83415 (United States)

    1999-01-01

    A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10{percent} of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon. {copyright} {ital 1999 American Institute of Physics.}

  15. Rocket Flight Path

    Directory of Open Access Journals (Sweden)

    Jamie Waters

    2014-09-01

    Full Text Available This project uses Newton’s Second Law of Motion, Euler’s method, basic physics, and basic calculus to model the flight path of a rocket. From this, one can find the height and velocity at any point from launch to the maximum altitude, or apogee. This can then be compared to the actual values to see if the method of estimation is a plausible. The rocket used for this project is modeled after Bullistic-1 which was launched by the Society of Aeronautics and Rocketry at the University of South Florida.

  16. Application of theory to propeller design

    Science.gov (United States)

    Cox, G. G.; Morgan, W. B.

    1974-01-01

    The various theories concerning propeller design are discussed. The use of digital computers to obtain specific blade shapes to meet appropriate flow conditions is emphasized. The development of lifting-line and lifting surface configurations is analyzed. Ship propulsive performance and basic propeller design considerations are investigated. The characteristics of supercavitating propellers are compared with those of subcavitating propellers.

  17. Preliminary Thermo-hydraulic Core Design Analysis of Korea Advanced Nuclear Thermal Engine Rocket for Space Application

    Energy Technology Data Exchange (ETDEWEB)

    Nam, Seung Hyun; Lee, Jeong Ik; Chang, Soon Heung [Korea Advanced Institute of Science and Technology, Daejeon (Korea, Republic of)

    2014-05-15

    Nclear rockets improve the propellant efficiency more than twice compared to CRs and thus significantly reduce the propellant requirement. The superior efficiency of nuclear rockets is due to the combination of the huge energy density and a single low molecular weight propellant utilization. Nuclear Thermal Rockets (NTRs) are particularly suitable for manned missions to Mars because it satisfies a relatively high thrust as well as a high propellant efficiency. NTRs use thermal energy released from a nuclear fission reactor to heat a single low molecular weight propellant, i. e., Hydrogen (H{sub 2}) and then exhausted the extremely heated propellant through a thermodynamic nozzle to produce thrust. A propellant efficiency parameter of rocket engines is specific impulse (I{sub sp}) which represents the ratio of the thrust over the rate of propellant consumption. The difference of I{sub sp} makes over three times propellant savings of NTRs for a manned Mars mission compared to CRs. NTRs can also be configured to operate bimodally by converting the surplus nuclear energy to auxiliary electric power required for the operation of a spacecraft. Moreover, the concept and technology of NTRs are very simple, already proven, and safe. Thus, NTRs can be applied to various space missions such as solar system exploration, International Space Station (ISS) transport support, Near Earth Objects (NEOs) interception, etc. Nuclear propulsion is the most promising and viable option to achieve challenging deep space missions. Particularly, the attractions of a NTR include excellent thrust and propellant efficiency, bimodal capability, proven technology, and safe and reliable performance. The ROK has also begun the research for space nuclear systems as a volunteer of the international space race and a major world nuclear energy country. KANUTER is one of the advanced NTR engines currently under development at KAIST. This bimodal engine is operated in two modes of propulsion with 100 MW

  18. Cavitation simulation on marine propellers

    DEFF Research Database (Denmark)

    Shin, Keun Woo

    Cavitation on marine propellers causes thrust breakdown, noise, vibration and erosion. The increasing demand for high-efficiency propellers makes it difficult to avoid the occurrence of cavitation. Currently, practical analysis of propeller cavitation depends on cavitation tunnel test, empirical...... criteria and inviscid flow method, but a series of model test is costly and the other two methods have low accuracy. Nowadays, computational fluid dynamics by using a viscous flow solver is common for practical industrial applications in many disciplines. Cavitation models in viscous flow solvers have been...... developed in the last decade. They show the potential for the simulation of propeller cavitation with robustness, but they are still to be more proved for practical applications. In the present work, hydrodynamic and numerical characteristics of several cavitation models developed for a viscous flow solver...

  19. Cryogenic rocket engine development at Delft aerospace rocket engineering

    NARCIS (Netherlands)

    Wink, J; Hermsen, R.; Huijsman, R; Akkermans, C.; Denies, L.; Barreiro, F.; Schutte, A.; Cervone, A.; Zandbergen, B.T.C.

    2016-01-01

    This paper describes the current developments regarding cryogenic rocket engine technology at Delft Aerospace Rocket Engineering (DARE). DARE is a student society based at Delft University of Technology with the goal of being the first student group in the world to launch a rocket into space. After

  20. Long-term firing tests of the nozzles of rocket engines made on the basis of carbon composite materials

    Science.gov (United States)

    Gubertov, A. M.; Koshlakov, V. V.; Mironov, V. V.; Rubinskii, V. R.; Pashutov, A. V.; Antipov, Ye. A.; Bratukhin, N. A.; Volkov, N. N.; Volkova, L. I.; Tsatsuev, S. M.; Tlevtsezhev, V. V.

    2012-12-01

    The results of the experimental investigation of the physico-chemical processes of interaction, destruction, and ablation of carbon composite materials and oxidation-protective coatings of nozzles of liquid-propellant rocket engines are presented. The thermally-stressed state of the joint between the nozzle made of composite material and the metallic combustion chamber of the rocket engine under standard operating conditions have been analyzed.

  1. Development and Testing of a Green-Propellant Micro-Hybrid Thruster with Electrostatic Ignition

    Science.gov (United States)

    Whitmore, Stephen A.; Judson, Michael D.

    2012-01-01

    , requiring an energy input of 14,850 Joules for catalytic dissociation. The hydrocarbon-seeded micro-hybrid was also adapted as a non-pyrotechnic ignitor for a 900 N (200-lbf) thrust hybrid motor. The motor was successfully ignited 4 consecutive times with no hardware swaps or propellant additions. The amount of ABS seed material that can be fit into the injector cap is the only limit to the number of available repeat firings. This series of tests marks the first time a hybrid motor was ever ignited by other than a solid-propellant pyrotechnic charge or bi-propellant flame ignitor. Nitrous oxide hybrid motors are typically difficult to ignite and usually require multiple solid-propellant charges to initiate combustion, so this nonpyrotechnic ignition is a significant accomplishment. The controlled hydrocarbon-seeding approach is fundamentally different from all other green propellant solutions offered by the aerospace industry. Although the proposed system is more correctly a hybrid technology; the system retains all the simple features of a monopropellant design. To date no optimization study has been performed to identify the best grain geometry for electrostatic ignition. Fortunately, because the grain segments are fabricated using rapid-prototyping technology, changing the grain geometry is as simple as modifying the 3-D printer CAD-file. Vacuum Isp exceeding 270 seconds has been demonstrated (Ref v), a value significantly higher than those offered by competing green monopropellant options. The propellants of choice, N2O/GOX and ABS are 100% non-toxic, non-explosive, and environmentally benign. Because the inert oxidizer and fuel components are mixed only within the combustion chamber, the system retains the inherent safety of a hybrid rocket and can be piggy-backed as a secondary payload with no overall mission risk increase to the primary payload, an excellent characteristic for secondary launch systems.

  2. DESIGN EKONOMIS UNTUK PROPELLER KAPAL

    Directory of Open Access Journals (Sweden)

    Hartono Hartono

    2012-02-01

    Full Text Available Increase of the price of world oil pushs liner to conduct action is economic from every operational unit in it’sship armada. One other most dominant in usage of fuel is Main engine is working to turn around propeller asship actuator. On that account ship owner wants design propeller which is economic for it’s ship to canreduce usage of fuel of 20% when sailing.

  3. Thrust Deduction in Contrarotating Propellers

    Science.gov (United States)

    1974-11-01

    Numerical results are given for a MARAD high -speed containership. Some principal findings for the subject ship are: (1) good agreement between...betwet.-i formard and after propellers. and the spacing between propellers. Computational re-wIlts air presen ted for a NIAHAl) high -speed...ratio of the effective velocity to the volume mean velocity. The volume mecan %eiocity ratio is (defined as ’V 2 h X h where V~ /V is the

  4. Satellite Propellant Pump Research

    Science.gov (United States)

    Schneider, Steven J.; Veres, Joseph P.; Hah, Chunill; Nerone, Anthony L.; Cunningham, Cameron C.; Kraft, Thomas G.; Tavernelli, Paul F.; Fraser, Bryan

    2005-01-01

    NASA Glenn initiated a satellite propellant pump technology demonstration program. The goal was to demonstrate the technologies for a 60 percent efficient pump at 1 gpm flow rate and 500 psia pressure rise. The pump design and analysis used the in-house developed computer codes named PUMPA and HPUMP3D. The requirements lead to a 4-stage impeller type pump design with a tip diameter of 0.54 inches and a rotational speed of 57,000 rpm. Analyses indicated that flow cavitation was not a problem in the design. Since the flow was incompressible, the stages were identical. Only the 2-stage pump was designed, fabricated, assembled, and tested for demonstration. Water was selected as the surrogate fluid for hydrazine in this program. Complete mechanical design including stress and dynamic analyses were conducted. The pump was driven by an electric motor directly coupled to the impellers. Runs up to 57,000 rpm were conducted, where a pressure rise of 200 psia at a flow rate of 0.8 gpm was measured to validate the design effort.

  5. Hydrocarbon Rocket Technology Impact Forecasting

    Science.gov (United States)

    Stuber, Eric; Prasadh, Nishant; Edwards, Stephen; Mavris, Dimitri N.

    2012-01-01

    Ever since the Apollo program ended, the development of launch propulsion systems in the US has fallen drastically, with only two new booster engine developments, the SSME and the RS-68, occurring in the past few decades.1 In recent years, however, there has been an increased interest in pursuing more effective launch propulsion technologies in the U.S., exemplified by the NASA Office of the Chief Technologist s inclusion of Launch Propulsion Systems as the first technological area in the Space Technology Roadmaps2. One area of particular interest to both government agencies and commercial entities has been the development of hydrocarbon engines; NASA and the Air Force Research Lab3 have expressed interest in the use of hydrocarbon fuels for their respective SLS Booster and Reusable Booster System concepts, and two major commercially-developed launch vehicles SpaceX s Falcon 9 and Orbital Sciences Antares feature engines that use RP-1 kerosene fuel. Compared to engines powered by liquid hydrogen, hydrocarbon-fueled engines have a greater propellant density (usually resulting in a lighter overall engine), produce greater propulsive force, possess easier fuel handling and loading, and for reusable vehicle concepts can provide a shorter turnaround time between launches. These benefits suggest that a hydrocarbon-fueled launch vehicle would allow for a cheap and frequent means of access to space.1 However, the time and money required for the development of a new engine still presents a major challenge. Long and costly design, development, testing and evaluation (DDT&E) programs underscore the importance of identifying critical technologies and prioritizing investment efforts. Trade studies must be performed on engine concepts examining the affordability, operability, and reliability of each concept, and quantifying the impacts of proposed technologies. These studies can be performed through use of the Technology Impact Forecasting (TIF) method. The Technology Impact

  6. Rockets in World War I

    Science.gov (United States)

    2004-01-01

    World War I enlisted rockets once again for military purposes. French pilots rigged rockets to the wing struts of their airplanes and aimed them at enemy observation balloons filled with highly inflammable hydrogen.

  7. Nitrous oxide cooling in hybrid rocket nozzles

    Science.gov (United States)

    Lemieux, Patrick

    2010-02-01

    The Department of Mechanical Engineering at the California Polytechnic State University, San Luis Obispo, has developed an innovative program of experimental research and development on hybrid rocket motors (where the fuel and the oxidizer are in different phases prior to combustion). One project currently underway involves the development of aerospike nozzles for such motors. These nozzles, however, are even more susceptible to throat ablation than regular converging-diverging nozzles, due the nature of their flow expansion mechanism. This paper presents the result of a recent development project focused on reducing throat ablation in hybrid rocket motor nozzles. Although the method is specifically targeted at increasing the life and operating range of aerospike nozzles, this paper describes its proof-of-concept implementation on conventional nozzles. The method is based on a regenerative cooling mechanism that differs in practice from that used in liquid propellant motors. A series of experimental tests demonstrate that this new method is not only effective at reducing damage in the most ablative region of the nozzle, but that the nozzle can survive multiple test runs.

  8. Large Liquid Rocket Testing: Strategies and Challenges

    Science.gov (United States)

    Rahman, Shamim A.; Hebert, Bartt J.

    2005-01-01

    Rocket propulsion development is enabled by rigorous ground testing in order to mitigate the propulsion systems risks that are inherent in space flight. This is true for virtually all propulsive devices of a space vehicle including liquid and solid rocket propulsion, chemical and non-chemical propulsion, boost stage and in-space propulsion and so forth. In particular, large liquid rocket propulsion development and testing over the past five decades of human and robotic space flight has involved a combination of component-level testing and engine-level testing to first demonstrate that the propulsion devices were designed to meet the specified requirements for the Earth to Orbit launchers that they powered. This was followed by a vigorous test campaign to demonstrate the designed propulsion articles over the required operational envelope, and over robust margins, such that a sufficiently reliable propulsion system is delivered prior to first flight. It is possible that hundreds of tests, and on the order of a hundred thousand test seconds, are needed to achieve a high-reliability, flight-ready, liquid rocket engine system. This paper overviews aspects of earlier and recent experience of liquid rocket propulsion testing at NASA Stennis Space Center, where full scale flight engines and flight stages, as well as a significant amount of development testing has taken place in the past decade. The liquid rocket testing experience discussed includes testing of engine components (gas generators, preburners, thrust chambers, pumps, powerheads), as well as engine systems and complete stages. The number of tests, accumulated test seconds, and years of test stand occupancy needed to meet varying test objectives, will be selectively discussed and compared for the wide variety of ground test work that has been conducted at Stennis for subscale and full scale liquid rocket devices. Since rocket propulsion is a crucial long-lead element of any space system acquisition or

  9. Bubble capture by a propeller

    Science.gov (United States)

    Caillé, François; Clanet, Christophe; Magnaudet, Jacques

    2006-08-01

    A small air bubble (radius a) is injected in water (kinematic viscosity nu) in the vicinity (distance r_0) of a propeller (radius r_p, angular frequency omega). We study experimentally and theoretically the conditions under which the bubble can be ‘captured’, i.e. deviated from its vertical trajectory (imposed by gravity g) and moved toward the centre of the propeller (r {=} 0). We show that the capture frequency omega_{scriptsizecapt} follows the relationship [omega_{hboxriptsizeit capt}=left(frac{2ga^2}{9betanu r_p f(hboxRe_b)}right)left(frac{r_0}{r_p}right)^2(1+\\cos\\varphi_0),] where beta is a dimensionless parameter characterizing the propeller, f(Re_b) is an empirical correction to Stokes' drag law which accounts for finite-Reynolds-number effects and pi/2-varphi_0 is the angle between the axis of the propeller and the line between the centre of the propeller and the point where the bubble is injected. This law is found to be valid as long as the distance d between the propeller and the water surface is larger than 3r_0. For smaller distances, the capture frequency increases; using an image technique, we show how the above expression is modified by the presence of the surface.

  10. Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

    Science.gov (United States)

    Kubiak, Jonathan M.; Arnett, Lori A.

    2016-01-01

    The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and

  11. ROCKETS: Soar to Success

    Science.gov (United States)

    Brett, Christine E. W.; O'Merle, Mary Jane; White, Gene

    2017-01-01

    This article describes ROCKETS, an after-school program for at-risk youth, and how the university students became involved in this service-learning project. The article discusses the steps that were taken to start the program, what is being done to continue the program, and the challenges that faculty have faced. This program is an authentic…

  12. The Relativistic Rocket

    Science.gov (United States)

    Antippa, Adel F.

    2009-01-01

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful…

  13. This "Is" Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

  14. Rocket Combustion Chamber Coating

    Science.gov (United States)

    Holmes, Richard R. (Inventor); McKechnie, Timothy N. (Inventor)

    2001-01-01

    A coating with the ability to protect (1) the inside wall (i.e., lining) of a rocket engine combustion chamber and (2) parts of other apparatuses that utilize or are exposed to combustive or high temperature environments. The novelty of this invention lies in the manner a protective coating is embedded into the lining.

  15. Baking Soda and Vinegar Rockets

    Science.gov (United States)

    Claycomb, James R.; Zachary, Christopher; Tran, Quoc

    2009-01-01

    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the…

  16. Seal testing of large diameter rocket motors

    Science.gov (United States)

    Moore, N. B.; Hellums, John W.; Bechtel, Thomas; Kittredge, Ken; Crossfield, Craig

    1991-01-01

    This investigative program examines leakage testing of elastomeric O-ring seals for a solid rocket casing and provides direction towards an improved nondestructive postassembly test. It also details test equipment for the Space Shuttle systems solid rocket boosters (SRB). The results are useful to designers of hardware for pressure containment vessels which use O-ring seals. Using several subscale seal and groove configuration test fixtures equipped with either two or three O-ring seals in series, seal integrity is investigated with both a pressure decay and flowmeter methods. Both types of test equipment adequately detect the practical range of expected seal leak rates of 1 to 0.0001 sccs. The flowmeter leak test equipment appears to reduce testing time substantially. Limited seal leakage testing is performed on full-sized rocket motor segment seals, a pre-Challenger short stack, providing comparison of bore seals to test specimen bore and face seals. The conclusions are that seal reliability, verified via a performance pressure test, can be affected by temperature, quantity of grease, test pressure, and seal pressure load direction. Potential seal failure scenarios including contamination, seal damage, and sealing surface damage are discussed. Recommendations are made for an improved test procedure.

  17. Experimental investigation of atomization characteristics of swirling spray by ADN gelled propellant

    Science.gov (United States)

    Guan, Hao-Sen; Li, Guo-Xiu; Zhang, Nai-Yuan

    2018-03-01

    Due to the current global energy shortage and increasingly serious environmental issues, green propellants are attracting more attention. In particular, the ammonium dinitramide (ADN)-based monopropellant thruster is gaining world-wide attention as a green, non-polluting and high specific impulse propellant. Gel propellants combine the advantages of liquid and solid propellants, and are becoming popular in the field of spaceflight. In this paper, a swirling atomization experimental study was carried out using an ADN aqueous gel propellant under different injection pressures. A high-speed camera and a Malvern laser particle size analyzer were used to study the spray process. The flow coefficient, cone angle of swirl atomizing spray, breakup length of spray membrane, and droplet size distribution were analyzed. Furthermore, the effects of different injection pressures on the swirling atomization characteristics were studied.

  18. Regenerative Gas Dryer for In-Situ Propellant Production

    Science.gov (United States)

    Paz, Aaron

    2017-01-01

    Rocket propellant can be produced anywhere that water is found by splitting it into hydrogen and oxygen, potentially saving several tons of mass per mission and enabling the long term presence of humans in space beyond LEO. When water is split into hydrogen and oxygen, the gaseous products can be very humid (several thousand ppm). Propellant-grade gases need to be extremely dry before being converted into cryogenic liquids (less than 26 ppm water for grade B Oxygen). The primary objective of this project is to design, build and test a regenerative gas drying system that can take humid gas from a water electrolysis system and provide dry gas (less than 26ppm water) to the inlet of a liquefaction system for long durations. State of the art work in this area attempted to use vacuum as a means to regenerate desiccant, but it was observed that water would migrate to the dry zone without a sweep gas present to direct the desorbed vapor. Further work attempted to use CO2 as a sweep gas, but this resulted in a corrosive carbonic acid. In order for in-situ propellant production to work, we need a way to continuously dry humid gas that addresses these issues.

  19. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    Science.gov (United States)

    Greene, William D. (Inventor)

    2009-01-01

    The proposed technology is a rocket engine cycle utilizing as the propulsive fluid a low molecular weight, cryogenic fluid, typically liquid hydrogen, pressure driven, heated, and expelled through a nozzle to generate high velocity and high specific impulse discharge gas. The proposed technology feeds the propellant through the engine cycle without the use of a separate pressurization fluid and without the use of turbomachinery. Advantages of the proposed technology are found in those elements of state-of-the-art systems that it avoids. It does not require a separate pressurization fluid or a thick-walled primary propellant tank as is typically required for a classical pressure-fed system. Further, it does not require the acceptance of intrinsic reliability risks associated with the use of turbomachinery

  20. Thermohydraulic Design Analysis Modeling for Korea Advanced NUclear Thermal Engine Rocket for Space Application

    Energy Technology Data Exchange (ETDEWEB)

    Nam, Seung Hyun; Choi, Jae Young; Venneria, Paolo F.; Jeong, Yong Hoon; Chang, Soon Heung [KAIST, Daejeon (Korea, Republic of)

    2015-05-15

    Space exploration is a realistic and profitable goal for long-term humanity survival, although the harsh space environment imposes lots of severe challenges to space pioneers. To date, almost all space programs have relied upon Chemical Rockets (CRs) rating superior thrust level to transit from the Earth's surface to its orbit. However, CRs inherently have insurmountable barrier to carry out deep space missions beyond Earth's orbit due to its low propellant efficiency, and ensuing enormous propellant requirement and launch costs. Meanwhile, nuclear rockets typically offer at least two times the propellant efficiency of a CR and thus notably reduce the propellant demand. Particularly, a Nuclear Thermal Rocket (NTR) is a leading candidate for near-term manned missions to Mars and beyond because it satisfies a relatively high thrust as well as a high efficiency. The superior efficiency of NTRs is due to both high energy density of nuclear fuel and the low molecular weight propellant of Hydrogen (H{sub 2}) over the chemical reaction by-products. A NTR uses thermal energy released from a nuclear fission reactor to heat the H{sub 2} propellant and then exhausted the highly heated propellant through a propelling nozzle to produce thrust. A propellant efficiency parameter of rocket engines is specific impulse (I{sub s}p) which represents the ratio of the thrust over the propellant consumption rate. If the average exhaust H{sub 2} temperature of a NTR is around 3,000 K, the I{sub s}p can be achieved as high as 1,000 s as compared with only 450 - 500 s of the best CRs. For this reason, NTRs are favored for various space applications such as orbital tugs, lunar transports, and manned missions to Mars and beyond. The best known NTR development effort was conducted from 1955 to1974 under the ROVER and NERVA programs in the USA. These programs had successfully designed and tested many different reactors and engines. After these projects, the researches on NERVA derived

  1. Nanotechnology Propellant Health Monitoring Sensors; Success Through Multi-Stakeholder Interests

    Science.gov (United States)

    2014-11-01

    be added • Reader access to secure server means real time updates Propellant aging sensor Shock sensor Passive RFID tag RFID reader Polymer Aging...currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 1. REPORT DATE (DD-MM-YYYY) September 2014 2. REPORT TYPE...rocket motors assessing both the current and future state.---AFRL SBIR solicitation “ Access viable approaches to cost effective, reliable service life

  2. Experimental Performance of a Novel Trochoidal Propeller

    Science.gov (United States)

    Roesler, Bernard; Epps, Brenden

    2015-11-01

    In the quest for energy efficiency in marine transportation, a promising marine propulsor concept is the trochoidal propeller. We have designed and tested a novel trochoidal propeller using a sinusoidal blade pitch function. The main results presented are measurements of thrust and torque, as well as the calculated efficiency, for a range of advance coefficients. The experimental data show narrow 95% confidence bounds, demonstrating high accuracy and repeatability in the experimental methods. We compare our sinusoidal-pitch trochoidal propeller with prior cross-flow propellers, as well as a representative screw propeller. While the efficiency of our propeller exceeds that of the cycloidal-pitch trochoidal propeller, it is slightly lower than the efficiencies of the other propellers considered. We also present a theoretical model that can be used to further explore and optimize such trochoidal propellers, leading to new avenues for improvements in marine propulsion systems.

  3. Aspects of Propeller Developements for a Submarine

    DEFF Research Database (Denmark)

    Andersen, Poul; kappel, Jens Julius; Spangenberg, Eugen

    2009-01-01

    Design and development of propellers for submarines are in some ways different from propellers for surface vessels. The most important demand is low acoustic signature that has priority over propeller efficiency, and the submarine propeller must be optimized with respect to acoustics rather than...... efficiency. Moreover the operating conditions of a submarine propeller are quite different. These aspects are discussed as well as the weighing of the various propeller parameters against the design objectives. The noise generated by the propeller can be characterized as thrust noise due to the inhomogeneous...... wake field of the submarine, trailing-edge noise and noise caused by turbulence in the inflow. The items discussed are demonstrated in a case study where a propeller of the Kappel type was developed. Three stages of the development are presented, including a design of an 8-bladed propeller where...

  4. Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

    Directory of Open Access Journals (Sweden)

    Bruce Chehroudi

    2012-01-01

    Full Text Available Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME or Vulcain (Ariane 5, the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL, spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

  5. Combustion and Performance Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

    Science.gov (United States)

    Hulka, J. R.; Jones, G. W.

    2010-01-01

    Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations. This paper summarizes the analyses of combustion and performance as a follow-up to a paper published in the 2008 JANNAF/LPS meeting. Combustion stability analyses are presented in a separate paper. The current paper includes test and analysis results of coaxial element injectors using liquid oxygen and liquid methane or gaseous methane propellants. Several thrust chamber configurations have been modeled, including thrust chambers with multi-element swirl coax element injectors tested at the NASA MSFC, and a uni-element chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods

  6. Ceramic Matrix Composite Turbine Disk for Rocket Engines

    Science.gov (United States)

    Effinger, Mike; Genge, Gary; Kiser, Doug

    2000-01-01

    NASA has recently completed testing of a ceramic matrix composite (CMC), integrally bladed disk (blisk) for rocket engine turbopumps. The turbopump's main function is to bring propellants from the tank to the combustion chamber at optimal pressures, temperatures, and flow rates. Advantages realized by using CMC blisks are increases in safety by increasing temperature margins and decreasing costs by increasing turbopump performance. A multidisciplinary team, involving materials, design, structural analysis, nondestructive inspection government, academia, and industry experts, was formed to accomplish the 4.5 year effort. This article will review some of the background and accomplishments of the CMC Blisk Program relative to the benefits of this technology.

  7. Hypergolic bipropellant spray combustion and flow modelling in rocket engines

    Science.gov (United States)

    Larosiliere, Louis M.; Litchford, Ron J.; Jeng, San-Mou

    1990-01-01

    A predictive tool for hypergolic bipropellant spray combustion and flow evolution in small rocket combustion chambers is described. It encompasses a computational technique for the gas-phase governing equations, a discrete particle method for liquid bipropellant sprays, and constitutive models for combustion chemistry, interphase exchanges, and unlike impinging hypergolic spray interactions. Emphasis is placed on the phenomenological modeling of the hypergolic liquid bipropellant gasification processes. Sample computations with the N2H4-N2O4 propellant system are given in order to show some of the capabilities and inadequacies of this tool.

  8. Schlieren image velocimetry measurements in a rocket engine exhaust plume

    Science.gov (United States)

    Morales, Rudy; Peguero, Julio; Hargather, Michael

    2017-11-01

    Schlieren image velocimetry (SIV) measures velocity fields by tracking the motion of naturally-occurring turbulent flow features in a compressible flow. Here the technique is applied to measuring the exhaust velocity profile of a liquid rocket engine. The SIV measurements presented include discussion of visibility of structures, image pre-processing for structure visibility, and ability to process resulting images using commercial particle image velocimetry (PIV) codes. The small-scale liquid bipropellant rocket engine operates on nitrous oxide and ethanol as propellants. Predictions of the exhaust velocity are obtained through NASA CEA calculations and simple compressible flow relationships, which are compared against the measured SIV profiles. Analysis of shear layer turbulence along the exhaust plume edge is also presented.

  9. Advanced active health monitoring system of liquid rocket engines

    Science.gov (United States)

    Qing, Xinlin P.; Wu, Zhanjun; Beard, Shawn; Chang, Fu-Kuo

    2008-11-01

    An advanced SMART TAPE system has been developed for real-time in-situ monitoring and long term tracking of structural integrity of pressure vessels in liquid rocket engines. The practical implementation of the structural health monitoring (SHM) system including distributed sensor network, portable diagnostic hardware and dedicated data analysis software is addressed based on the harsh operating environment. Extensive tests were conducted on a simulated large booster LOX-H2 engine propellant duct to evaluate the survivability and functionality of the system under the operating conditions of typical liquid rocket engines such as cryogenic temperature, vibration loads. The test results demonstrated that the developed SHM system could survive the combined cryogenic temperature and vibration environments and effectively detect cracks as small as 2 mm.

  10. Coolant Design System for Liquid Propellant Aerospike Engines

    Science.gov (United States)

    McConnell, Miranda; Branam, Richard

    2015-11-01

    Liquid propellant rocket engines burn at incredibly high temperatures making it difficult to design an effective coolant system. These particular engines prove to be extremely useful by powering the rocket with a variable thrust that is ideal for space travel. When combined with aerospike engine nozzles, which provide maximum thrust efficiency, this class of rockets offers a promising future for rocketry. In order to troubleshoot the problems that high combustion chamber temperatures pose, this research took a computational approach to heat analysis. Chambers milled into the combustion chamber walls, lined by a copper cover, were tested for their efficiency in cooling the hot copper wall. Various aspect ratios and coolants were explored for the maximum wall temperature by developing our own MATLAB code. The code uses a nodal temperature analysis with conduction and convection equations and assumes no internal heat generation. This heat transfer research will show oxygen is a better coolant than water, and higher aspect ratios are less efficient at cooling. This project funded by NSF REU Grant 1358991.

  11. Experimental Study of Ballistic-Missile Base Heating with Operating Rocket

    Science.gov (United States)

    Nettle, J. Cary

    1958-01-01

    A rocket of the 1000-pound-thrust class using liquid oxygen and JP-4 fuel as propellant was installed in the Lewis 8- by 6-foot tunnel to permit a controlled study of some of the factors affecting the heating of a rocket-missile base. Temperatures measured in the base region are presented from findings of three motor extension lengths relative to the base. Data are also presented for two combustion efficiency levels in the rocket motor. Temperature as high as 1200 F was measured in the base region because of the ignition of burnable rocket gases. combustibles that are dumped into the base by accessories seriously aggravate the base-burning temperature rise.

  12. Single stage to orbit mass budgets derived from propellant density and specific impulse

    Energy Technology Data Exchange (ETDEWEB)

    Whitehead, J.C.

    1996-06-06

    The trade between specific impulse (Isp) and density is examined in view of Single Stage To Orbit (SSTO) requirements. Mass allocations for vehicle hardware are derived from these two properties, for several propellant combinations and a dual-fuel case. This comparative analysis, based on flight-proven hardware, indicates that the higher density of several alternative propellants compensates for reduced Isp, when compared with cryogenic oxygen and hydrogen. Approximately half the orbiting mass of a rocket- propelled SSTO vehicle must be allocated to propulsion hardware and residuals. Using hydrogen as the only fuel requires a slightly greater fraction of orbiting mass for propulsion, because hydrogen engines and tanks are heavier than those for denser fuels. The advantage of burning both a dense fuel and hydrogen in succession depends strongly on tripropellant engine weight. The implications of the calculations for SSTO vehicle design are discussed, especially with regard to the necessity to minimize non-tankage structure.

  13. Thermal decomposition of AP/HTPB propellants in presence of Zn nanoalloys

    Science.gov (United States)

    Chaturvedi, Shalini; Dave, Pragnesh N.; Patel, Nikul N.

    2015-01-01

    Composite solid propellants were prepared with and without nanoalloys (Zn-Cu, Zn-Ni, Zn-Fe), where nanoalloys are used as catalyst. Catalytic properties of these nanomaterials measured on ammonium perchlorate/hydroxyl-terminated polybutadiene propellant by thermogravimetric analysis and differential thermal analysis. Both experimental results show enhancement in the thermal decomposition of propellants in presence of nanoalloys. In differential thermal analysis method, experiments had done at three heating rates, β1 = 5°, β2 = 10°, β3 = 15° per minute. Calculation of activation energy of high temperature decomposition step was done by using following Kissinger equation. Zn-Cu was found to be the best.

  14. The relativistic rocket

    Energy Technology Data Exchange (ETDEWEB)

    Antippa, Adel F [Departement de Physique, Universite du Quebec a Trois-Rivieres, Trois-Rivieres, Quebec G9A 5H7 (Canada)

    2009-05-15

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful method that can be applied to a wide range of special relativistic problems of linear acceleration.

  15. Multivariable optimization of liquid rocket engines using particle swarm algorithms

    Science.gov (United States)

    Jones, Daniel Ray

    Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

  16. Marine propellers: the latest topics.

    Science.gov (United States)

    Kubo, H

    1996-02-01

    The impeller of the axial flow blood pump in an artificial heart is essentially based on the same principle as a marine propeller. Impellers designed for artificial hearts and marine propellers have a number of points in common. Decreased cavitation and relieved fluctuation load are only representative of them. As for a distinct concept of pressure distribution, the inverse method could be very useful. Skew may led to a more mild and natural character in the blood. Highly skewed blades and super elastic blades have the potential to decrease the burden on the entire circulatory system. This paper will address the main points and latest issues in propeller design concluding with a discussion of the implications of these issues for blood pump impellers.

  17. A History of Collapse Factor Modeling and Empirical Data for Cryogenic Propellant Tanks

    Science.gov (United States)

    deQuay, Laurence; Hodge, B. Keith

    2010-01-01

    One of the major technical problems associated with cryogenic liquid propellant systems used to supply rocket engines and their subassemblies and components is the phenomenon of propellant tank pressurant and ullage gas collapse. This collapse is mainly caused by heat transfer from ullage gas to tank walls and interfacing propellant, which are both at temperatures well below those of this gas. Mass transfer between ullage gas and cryogenic propellant can also occur and have minor to significant secondary effects that can increase or decrease ullage gas collapse. Pressurant gas is supplied into cryogenic propellant tanks in order to initially pressurize these tanks and then maintain required pressures as propellant is expelled from these tanks. The net effect of pressurant and ullage gas collapse is increased total mass and mass flow rate requirements of pressurant gases. For flight vehicles this leads to significant and undesirable weight penalties. For rocket engine component and subassembly ground test facilities this results in significantly increased facility hardware, construction, and operational costs. "Collapse Factor" is a parameter used to quantify the pressurant and ullage gas collapse. Accurate prediction of collapse factors, through analytical methods and modeling tools, and collection and evaluation of collapse factor data has evolved over the years since the start of space exploration programs in the 1950 s. Through the years, numerous documents have been published to preserve results of studies associated with the collapse factor phenomenon. This paper presents a summary and selected details of prior literature that document the aforementioned studies. Additionally other literature that present studies and results of heat and mass transfer processes, related to or providing important insights or analytical methods for the studies of collapse factor, are presented.

  18. [Characteristics and mechanism of boat propeller injuries].

    Science.gov (United States)

    Yu, Song; Shen, Yi-Wen; Xue, Ai-Min

    2008-02-01

    To summarize the characteristics and investigate the mechanisms of boat propeller injuries so as to explore the identification methods between boat propeller injuries and corpse dismemberment. More than 100 autopsy cases of boat propeller injuries were collected in a period between 1994 and 2005 in Huzhou district, Zhejiang province. The characteristics of injuries caused by propeller, including abrasion, wound, fracture and severed wound, and the characteristics of clothing, were retrospectively studied and summarized. The severed cross wound section of boat propeller injuries was compared with that caused by corpse dismemberment. The boat propeller injuries were resulted from high-speed propellers with enormous splitting power and mechanical cutting, while corpse dismemberment were resulted from cutting and dismembering the body with sharp instruments. Due to the different mechanisms, the different strength of force and recoil force, the severed wound cross section had different characteristics. Wounds caused by boat propeller injuries have their unique characteristics, distinguished from wounds of dismembered corpse.

  19. Quadcopter thrust optimization with ducted-propeller

    Directory of Open Access Journals (Sweden)

    Kuantama Endrowednes

    2017-01-01

    Full Text Available In relation to quadcopter body frame model, propeller can be categorized into propeller with ducted and without ducted. This study present differences between those two using CFD (Computational Fluid Dynamics method. Both categories utilize two blade-propeller with diameter of 406 (mm. Propeller rotation generates acceleration per time unit on the volume of air. Based on the behavior of generated air velocity, ducted propeller can be modeled into three versions. The generated thrust and performance on each model were calculated to determine the best model. The use of ducted propeller increases the total weight of quadcopter and also total thrust. The influence of this modeling were analyzed in detail with variation of angular velocity propeller from 1000 (rpm to 9000 (rpm. Besides the distance between propeller tip and ducted barrier, the size of ducted is also an important part in thrust optimization and total weight minimization of quadcopter.

  20. Rocket-Powered Parachutes Rescue Entire Planes

    Science.gov (United States)

    2010-01-01

    Small Business Innovation Research (SBIR) contracts with Langley Research Center helped BRS Aerospace, of Saint Paul, Minnesota, to develop technology that has saved 246 lives to date. The company s whole aircraft parachute systems deploy in less than 1 second thanks to solid rocket motors and are capable of arresting the descent of a small aircraft, lowering it safely to the ground. BRS has sold more than 30,000 systems worldwide, and the technology is now standard equipment on many of the world s top-selling aircraft. Parachutes for larger airplanes are in the works.

  1. Plasma ignition of LOVA propellants

    NARCIS (Netherlands)

    Driel, C.A. van; Boluijt, A.G.; Schilt, A.

    2010-01-01

    Ignition experiments were performed using a gun simulator which is equipped with a burst disk. This equipment facilitates the application of propellant loading densities which are comparable to those applied in regular ammunitions. For this study the gun simulator was equipped with a plasma jet

  2. Propelling arboriculture into the future

    Science.gov (United States)

    E. Gregory McPherson

    2011-01-01

    Research is the engine that propels arboriculture and urban forestry into the future. New knowledge, technologies, and tools provide arborists with improved tree care practices that result in healthier urban forests. The ISA Science and Research Committee (SRC) is composed of 13 professionals and researchers who are dedicated to elevating the importance of research...

  3. Environmentally Friendly Advanced Gun Propellants

    Science.gov (United States)

    2004-09-15

    yellow and had the consistency of cookie dough . Total mix cycle was 90 to 120 minutes. Quarter-pint mixes were 80 grams and one-pint mixes were 275...Elastomer Laminated Propellant Development,” 34th JANNAF Combustion Subcommittee and Propulsion Systems Hazards Subcommittee and Airbreathing

  4. Cars Spectroscopy of Propellant Flames

    Science.gov (United States)

    1983-11-01

    Propellant Flames," Fast Reactions in Energetic Systems, D. Capellos and R. F. Walker, ed., Reidel, Boston, MA, 1981, pp 473-434. 2. L. E. Harris and M. E...Beardell Y. Carignon J. Fendell K, Aron E. Petro DRStfC-LCE-(D), R. Walker P. Marinkas C. Capellos S. Buluou F. Gilbert Dover, tU 07801 Afmtnistrator

  5. Innovative boron nitride-doped propellants

    Directory of Open Access Journals (Sweden)

    Thelma Manning

    2016-04-01

    Full Text Available The U.S. military has a need for more powerful propellants with balanced/stoichiometric amounts of fuel and oxidants. However, balanced and more powerful propellants lead to accelerated gun barrel erosion and markedly shortened useful barrel life. Boron nitride (BN is an interesting potential additive for propellants that could reduce gun wear effects in advanced propellants (US patent pending 2015-026P. Hexagonal boron nitride is a good lubricant that can provide wear resistance and lower flame temperatures for gun barrels. Further, boron can dope steel, which drastically improves its strength and wear resistance, and can block the formation of softer carbides. A scalable synthesis method for producing boron nitride nano-particles that can be readily dispersed into propellants has been developed. Even dispersion of the nano-particles in a double-base propellant has been demonstrated using a solvent-based processing approach. Stability of a composite propellant with the BN additive was verified. In this paper, results from propellant testing of boron nitride nano-composite propellants are presented, including closed bomb and wear and erosion testing. Detailed characterization of the erosion tester substrates before and after firing was obtained by electron microscopy, inductively coupled plasma and x-ray photoelectron spectroscopy. This promising boron nitride additive shows the ability to improve gun wear and erosion resistance without any destabilizing effects to the propellant. Potential applications could include less erosive propellants in propellant ammunition for large, medium and small diameter fire arms.

  6. Rocket + Science = Dialogue

    Science.gov (United States)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

    It's a cliche that rocket engineers and space scientists don t see eye-to-eye. That goes double for rocket engineers working on human spaceflight and scientists working on space telescopes and planetary probes. They work fundamentally different problems but often feel that they are competing for the same pot of money. Put the two groups together for a weekend, and the results could be unscientific or perhaps combustible. Fortunately, that wasn't the case when NASA put heavy lift launch vehicle designers together with astronomers and planetary scientists for two weekend workshops in 2008. The goal was to bring the top people from both groups together to see how the mass and volume capabilities of NASA's Ares V heavy lift launch vehicle could benefit the science community. Ares V is part of NASA's Constellation Program for resuming human exploration beyond low Earth orbit, starting with missions to the Moon. In the current mission scenario, Ares V launches a lunar lander into Earth orbit. A smaller Ares I rocket launches the Orion crew vehicle with up to four astronauts. Orion docks with the lander, attached to the Ares V Earth departure stage. The stage fires its engine to send the mated spacecraft to the Moon. Standing 360 feet high and weighing 7.4 million pounds, NASA's new heavy lifter will be bigger than the 1960s-era Saturn V. It can launch almost 60 percent more payload to translunar insertion together with the Ares I and 35 percent more mass to low Earth orbit than the Saturn V. This super-sized capability is, in short, designed to send more people to more places to do more things than the six Apollo missions.

  7. Liquid Bismuth Propellant Flow Sensor

    Science.gov (United States)

    Polzin, Kurt A.; Stanojev, B. J.; Korman, V.

    2007-01-01

    Quantifying the propellant mass flow rate in liquid bismuth-fed electric propulsion systems has two challenging facets. First, the flow sensors must be capable of providing a resolvable measurement at propellant mass flow rates on the order of 10 mg/see with and uncertainty of less that 5%. The second challenge has to do with the fact that the materials from which the flow sensors are fabricated must be capable of resisting any of the corrosive effects associated with the high-temperature propellant. The measurement itself is necessary in order to properly assess the performance (thrust efficiency, Isp) of thruster systems in the laboratory environment. The hotspot sensor[I] has been designed to provide the bismuth propellant mass flow rate measurement. In the hotspot sensor, a pulse of thermal energy (derived from a current pulse and associated joule heating) is applied near the inlet of the sensor. The flow is "tagged" with a thermal feature that is convected downstream by the flowing liquid metal. Downstream, a temperature measurement is performed to detect a "ripple" in the local temperature associated with the passing "hotspot" in the propellant. By measuring the time between the upstream generation and downstream detection of the thermal feature, the flow speed can be calculated using a "time of flight" analysis. In addition, the system can be calibrated by measuring the accumulated mass exiting the system as a-function of time and correlating this with the time it takes the hotspot to convect through the sensor. The primary advantage of this technique is that it doesn't depend on an absolute measurement of temperature but, instead, relies on the observation of thermal features. This makes the technique insensitive to other externally generated thermal fluctuations. In this paper, we describe experiments performed using the hotspot flow sensor aimed at quantifying the resolution of the sensor technology. Propellant is expelled onto an electronic scale to

  8. CFD Analysis of Square Flow Channel in Thermal Engine Rocket Adventurer for Space Nuclear Application

    Energy Technology Data Exchange (ETDEWEB)

    Nam, S. H.; Suh, K. Y. [Seoul National University, Seoul (Korea, Republic of); Kang, S. G. [PHILOSOPHIA, Inc., Seoul (Korea, Republic of)

    2008-10-15

    Solar system exploration relying on chemical rockets suffers from long trip time and high cost. In this regard nuclear propulsion is an attractive option for space exploration. The performance of Nuclear Thermal Rocket (NTR) is more than twice that of the best chemical rocket. Resorting to the pure hydrogen (H{sub 2}) propellant the NTRs can possibly achieve as high as 1,000 s of specific impulse (I{sub sp}) representing the ratio of the thrust over the fuel consumption rate, as compared to only 425 s of H{sub 2}/O{sub 2} rockets. If we reflect on the mission to Mars, NTRs would reduce the round trip time to less than 300 days, instead of over 600 days with chemical rockets. This work presents CFD analysis of one Fuel Element (FE) of Thermal Engine Rocket Adventurer (TERA). In particular, one Square Flow Channel (SFC) is analyzed in Square Lattice Honeycomb (SLHC) fuel to examine the effects of mass flow rate on rocket performance.

  9. Rocket Motor Joint Construction Including Thermal Barrier

    Science.gov (United States)

    Steinetz, Bruce M. (Inventor); Dunlap, Patrick H., Jr. (Inventor)

    2002-01-01

    A thermal barrier for extremely high temperature applications consists of a carbon fiber core and one or more layers of braided carbon fibers surrounding the core. The thermal barrier is preferably a large diameter ring, having a relatively small cross-section. The thermal barrier is particularly suited for use as part of a joint structure in solid rocket motor casings to protect low temperature elements such as the primary and secondary elastomeric O-ring seals therein from high temperature gases of the rocket motor. The thermal barrier exhibits adequate porosity to allow pressure to reach the radially outward disposed O-ring seals allowing them to seat and perform the primary sealing function. The thermal barrier is disposed in a cavity or groove in the casing joint, between the hot propulsion gases interior of the rocket motor and primary and secondary O-ring seals. The characteristics of the thermal barrier may be enhanced in different applications by the inclusion of certain compounds in the casing joint, by the inclusion of RTV sealant or similar materials at the site of the thermal barrier, and/or by the incorporation of a metal core or plurality of metal braids within the carbon braid in the thermal barrier structure.

  10. Rocket Assembly and Checkout Facility

    Data.gov (United States)

    Federal Laboratory Consortium — FUNCTION: Integrates, tests, and calibrates scientific instruments flown on sounding rocket payloads. The scientific instruments are assembled on an optical bench;...

  11. 14 CFR 35.23 - Propeller control system.

    Science.gov (United States)

    2010-01-01

    ... hazardous propeller effect. (3) The loss of normal propeller pitch control does not cause a hazardous propeller effect under the intended operating conditions. (4) The failure or corruption of data or signals shared across propellers does not cause a hazardous propeller effect. (c) Electronic propeller control...

  12. Rhenium Rocket Manufacturing Technology

    Science.gov (United States)

    1997-01-01

    The NASA Lewis Research Center's On-Board Propulsion Branch has a research and technology program to develop high-temperature (2200 C), iridium-coated rhenium rocket chamber materials for radiation-cooled rockets in satellite propulsion systems. Although successful material demonstrations have gained much industry interest, acceptance of the technology has been hindered by a lack of demonstrated joining technologies and a sparse materials property data base. To alleviate these concerns, we fabricated rhenium to C-103 alloy joints by three methods: explosive bonding, diffusion bonding, and brazing. The joints were tested by simulating their incorporation into a structure by welding and by simulating high-temperature operation. Test results show that the shear strength of the joints degrades with welding and elevated temperature operation but that it is adequate for the application. Rhenium is known to form brittle intermetallics with a number of elements, and this phenomena is suspected to cause the strength degradation. Further bonding tests with a tantalum diffusion barrier between the rhenium and C-103 is planned to prevent the formation of brittle intermetallics.

  13. Drag and Torque on Locked Screw Propeller

    Directory of Open Access Journals (Sweden)

    Tomasz Tabaczek

    2014-09-01

    Full Text Available Few data on drag and torque on locked propeller towed in water are available in literature. Those data refer to propellers of specific geometry (number of blades, blade area, pitch and skew of blades. The estimation of drag and torque of an arbitrary propeller considered in analysis of ship resistance or propulsion is laborious. The authors collected and reviewed test data available in the literature. Based on collected data there were developed the empirical formulae for estimation of hydrodynamic drag and torque acting on locked screw propeller. Supplementary CFD computations were carried out in order to prove the applicability of the formulae to modern moderately skewed screw propellers.

  14. The open-cycle gas-core nuclear rocket engine - Some engineering considerations.

    Science.gov (United States)

    Taylor, M. F.; Whitmarsh, C. L., Jr.; Sirocky, P. J., Jr.; Iwanczyk, L. C.

    1971-01-01

    A preliminary design study of a conceptual 6000-MW open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 44,200 lb and a specific impulse of 4400 sec. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel) and the waste heat rejection system were considered conceptually and were sized.

  15. Design and numerical investigation of swirl recovery vanes for the Fokker 29 propeller

    Directory of Open Access Journals (Sweden)

    Wang Yangang

    2014-10-01

    Full Text Available Swirl recovery vanes (SRVs are a set of stationary vanes located downstream from a propeller, which may recover some of the residual swirl from the propeller, hoping for an improvement in both thrust and efficiency. The SRV concept design for a scaled version representing the Fokker 29 propeller is performed in this paper, which may give rise to a promotion in propulsive performance of this traditional propeller. Firstly the numerical strategy is validated from two aspects of global quantities and the local flow field of the propeller compared with experimental data, and then the exit flow together with the development of propeller wake is analyzed in detail. Three kinds of SRV are designed with multiple circular airfoils. The numerical results show that the swirl behind the propeller is recovered significantly with Model V3, which is characterized by the highest solidity along spanwise, for various working conditions, and the combination of rotor and vane produced 5.76% extra thrust at the design point. However, a lower efficiency is observed asking for a better vane design and the choice of a working point. The vane position is studied which shows that there is an optimum range for higher thrust and efficiency.

  16. What fuel for a rocket?

    CERN Document Server

    Miranda, E N

    2012-01-01

    Elementary concepts from general physics and thermodynamics have been used to analyze rocket propulsion. Making some reasonable assumptions, an expression for the exit velocity of the gases is found. From that expression one can conclude what are the desired properties for a rocket fuel.

  17. Development Testing of 1-Newton ADN-Based Rocket Engines

    Science.gov (United States)

    Anflo, K.; Gronland, T.-A.; Bergman, G.; Nedar, R.; Thormählen, P.

    2004-10-01

    With the objective to reduce operational hazards and improve specific and density impulse as compared with hydrazine, the Research and Development (R&D) of a new monopropellant for space applications based on AmmoniumDiNitramide (ADN), was first proposed in 1997. This pioneering work has been described in previous papers1,2,3,4 . From the discussion above, it is clear that cost savings as well as risk reduction are the main drivers to develop a new generation of reduced hazard propellants. However, this alone is not enough to convince a spacecraft builder to choose a new technology. Cost, risk and schedule reduction are good incentives, but a spacecraft supplier will ask for evidence that this new propulsion system meets a number of requirements within the following areas: This paper describes the ongoing effort to develop a storable liquid monopropellant blend, based on AND, and its specific rocket engines. After building and testing more than 20 experimental rocket engines, the first Engineering Model (EM-1) has now accumulated more than 1 hour of firing-time. The results from test firings have validated the design. Specific impulse, combustion stability, blow-down capability and short pulse capability are amongst the requirements that have been demonstrated. The LMP-103x propellant candidate has been stored for more than 1 year and initial material compatibility screening and testing has started. 1. Performance &life 2. Impact on spacecraft design &operation 3. Flight heritage Hereafter, the essential requirements for some of these areas are outlined. These issues are discussed in detail in a previous paper1 . The use of "Commercial Of The Shelf" (COTS) propulsion system components as much as possible is essential to minimize the overall cost, risk and schedule. This leads to the conclusion that the Technology Readiness Level (TRL) 5 has been reached for the thruster and propellant. Furthermore, that the concept of ADN-based propulsion is feasible.

  18. Thermal Degradation Studies of A Polyurethane Propellant Binder

    Energy Technology Data Exchange (ETDEWEB)

    Assink, R.A.; Celina, M.; Gillen, K.T.; Graham, A.C.; Minier, L.M.

    1999-06-12

    The thermal oxidative aging of a crosslinked hydroxy-terminated polybutadiene (HTPB)/isophorone diisocyanate (IPDI) based polyurethane rubber, used as a polymeric binder in solid propellant grain, was investigated at temperatures from 25 C to 125 C. The changes in tensile elongation, polymer network properties and chain dynamics, mechanical hardening and density were determined with a range of techniques including modulus profiling, solvent swelling, NMR relaxation and O{sub 2} permeability measurements. We critically evaluated the Arrhenius methodology that is commonly used with a linear extrapolation of high temperature aging data using extensive data superposition and highly sensitive oxygen consumption experiments. The effects of other constituents in the propellant formulation on aging were also investigated. We conclude that crosslinking is the dominant process at higher temperatures and that the degradation involves only limited hardening in the bulk of the material. Significant curvature in the Arrhenius diagram of the oxidation rates was observed. This is similar to results for other rubber materials.

  19. Thermal decomposition of phase-stabilised ammonium nitrate (PSAM), HTPB based propellants. The effect of iron(III)oxide burning-rate catalyst

    NARCIS (Netherlands)

    Carvalheira, P.; Gadiot, G.M.H.J.L.; Klerk, W.P.C. de

    1995-01-01

    Phase-stabilised ammonium nitrate (PSAN) and hydroxyl-terminated polybutadiene (HTPB) are the main ingredients of propellants used with success in some pyrotechnic igniter components of the VULCAIN liquid rocket engine for the ARIANE 5. Small amounts of selected additives play an important role in

  20. Nuclear Rocket Engine Reactor

    CERN Document Server

    Lanin, Anatoly

    2013-01-01

    The development of a nuclear rocket engine reactor (NRER ) is presented in this book. The working capacity of an active zone NRER under mechanical and thermal load, intensive neutron fluxes, high energy generation (up to 30 MBT/l) in a working medium (hydrogen) at temperatures up to 3100 K is displayed. Design principles and bearing capacity of reactors area discussed on the basis of simulation experiments and test data of a prototype reactor. Property data of dense constructional, porous thermal insulating and fuel materials like carbide and uranium carbide compounds in the temperatures interval 300 - 3000 K are presented. Technological aspects of strength and thermal strength resistance of materials are considered. The design procedure of possible emergency processes in the NRER is developed and risks for their origination are evaluated. Prospects of the NRER development for pilotless space devices and piloted interplanetary ships are viewed.